[1st Edition] 9781483224664

Table of contents :
Content:
Front MatterPage iii
Copyright pagePage iv
Contributors to Volume 5Page v
ForewordPages vii-xFREDERICK I. ORDWAY III
Contents of Previous VolumesPages xv-xvi
Supplementary MonographPage xvii
Astronautical Investigations of the SunPages 1-20R. GRANT ATHAY
Advances in Communication Relay Satellite TechniquesPages 21-46R.P. HAVILAND
Solid Propellant Rocket TechnologyPages 47-86H.W. RITCHEY, J.M. MCDERMOTT
Environmental Control of Manned Space VehiclesPages 87-142ROBERT E. SMITH
Terrestrial, Lunar, and Planetary Applications of Navigation and Geodetic SatellitesPages 143-230JOHN D. NICOLAIDES, MARK M. MACOMBER, WILLIAM M. KAULA
Orbital OperationsPages 231-325KRAFFT A. EHRICKE
Author IndexPages 327-331
Subject IndexPages 332-334

Citation preview

Advances in

Space Science and Technology Edited

by

FREDERICK I. ORDWAY, III General Astronautics Research Corporation Washington, D. C. and Space Science and Technology Information Center Huntsville, Alabama

Editorial Advisory Board

Wernher von Braun Frederick C. Durant, III Hideo Itokawa David F. Martyn Eugen Sanger Leslie R. Shepherd George P. Sutton Etienne Vassy

VOLUME 5

ACADEMIC PRESS

NEW YORK and LONDON

1963

COPYRIGHT © 1 9 6 3 , BY ACADEMIC PRESS I N C . ALL RIGHTS RESERVED. NO PART OF THIS BOOK MAY BE REPRODUCED IN ANY FORM, BY PHOTOSTAT, MICROFILM, OR ANY OTHER MEANS, WITHOUT WRITTEN PERMISSION FROM THE PUBLISHERS.

ACADEMIC PRESS INC. I l l Fifth Avenue, New York 3 , New York

United Kingdom Edition published by ACADEMIC PRESS INC. (LONDON) LTD. Berkeley Square House, London W.l

LIBRARY OF CONGRESS CATALOG CARD NUMBER: 5 9 - 1 5 7 6 0

PRINTED I N THE UNITED STATES OF AMERICA

Contributors to Volume 5 Department of Astrophysics and Atmospheric of Colorado, and High Altitude Observatory,

R . GRANT ATHAY,

University Colorado

Physics, Boulder,

Advanced Studies Office, General Dynamics/Astro­ nautics, San Diego, California R . P . HAVILAND, Missile and Space Division, General Electric Company, Philadelphia, Pennsylvania WILLIAM M . KAULA, Goddard Space Flight Center, National Aeronautics and Space Administration, Greenbelt, Maryland MARK M . MACOMBER, Bureau of Naval Weapons, Department of the Navy, Washington, D. C. J . M . MCDERMOTT, Development Laboratories, Thiokol Chemical Corpo­ ration, Wasatch Division, Brigham City, Utah JOHN D . NICOLAIDES, Office of Space Sciences, National Aeronautics and Space Administration, Washington, D. C. H . W . RITCHEY, Thiokol Chemical Corporation, Rocket Operations Center, Ogden, Utah ROBERT E . SMITH, Department of Physiology, School of Medicine, Uni­ versity of California Medical Center, Los Angeles, California KRAFFT A . EHRICKE,

v

Foreword Six chapters make up this volume of Advances in Space Science and Technology, the fifth in the series. A seventh chapter was expanded into a monograph, becoming Supplement No. 1. It is entitled "Space Carrier Vehicles," is by Dr. Oswald H. Lange and Richard J. Stein, and treats the design, development, and testing of launching rockets. In the intervening year since the publication of Volume 4 the astronautical community has settled down to the sustained, hard, and often tedius task of developing truly reliable, long-lifetime, manned space vehicles whose medium-term objective is the landing of men on the surface of the Moon. In mid-May of 1963 the Mercury manned satellite program came to a close with the orbiting of the 3,033-lb MA9 capsule with astronaut Leroy Gordon Cooper aboard. A month later the Soviets launched Vostoks 5 and 6 two days apart with Lt. Col. Valery F. Bykovsky aboard the former and Miss Valentina V. Tereshkova aboard the latter; they spent 119 hr 4 min and 70 hr 40 min in space, respectively. Progress on the U. S. two-man Gemini satellite was recorded, and signifi­ cant milestones were passed on the three-man Apollo lunar spaceship program. The fate of the Dynasoar orbital research remained unclear as did the entire military position in space. There were no significant events in lunar and planetary exploration. Following the failure of Ranger 5 to meet objectives in mid-October 1962 the program was reevaluated and firings postponed. The Soviet Automaticheskaya Mezhplanetnaya Stantsiya 3 was fired toward Mars in November 1962, but toward the end of March 1963 radio contact was lost. Automaticheskaya Mezhplanetnaya Stantsiya 4 by-passed the Moon following launch on 2 April 1963 and entered into an elongated Earth orbit. Few details were given concerning the scientific achievements of the flight. It was speculated that it was intended to soft land on the Moon but for some reason failed. As in previous years, a series of unmanned satellites was successfully launched into Earth orbit. Some satellites were military and hence in­ formation concerning them is scarce. Others, like the Soviet Kosmos series, were ostensibly scientific but just as few data became available on their vii

viii

FOREWORD

purposes and achievements. And others were widely publicized, the U. S. communication satellites Relay 1, Syncom 1 and 2, and Telstar 2 being the best examples. The use of satellites to relay television programs across the Atlantic was almost considered commonplace as the world became increasingly accustomed to, and to an extent decreasingly interested in, the marvels of astronautics. While firings of instrumented vehicles into the reaches of the Solar System are still comparatively rare, ways and means of sending them on their long and lonely voyages are steadily being perfected. In earlier volumes of this series authorities wrote on the astronautical investigation of the Moon and of the planets and their satellites. Now, in the present volume, a recognized expert turns his attention to the Sun. In successive sections he (1) looks into the major unsolved solar problems (thermodynamic structure, geometrical structure, velocity fields, flares and other transient phenomena, solar magnetic fields, and corpus­ cular emission), (2) analyzes the role of space observations in solar physics, and (3) considers the demands placed on laboratory physics. We are cautioned that "for every problem related to the Sun that we claim to T understand, w e can name many related problems that at best are only poorly understood." And, furthermore, we must be prepared to accept the fact that "there is much happening on the Sun of which we are still totally unaware." It is recommended that we consider the space exploration of the Sun as "a means of pioneering into new realms of solar physics, both by searching out suspected phenomena that are shielded from ground-based observations by our terrestrial atmosphere and by seeking out new phenomena as a means of broadening the scope of our research efforts in both astrophysics and geophysics." The second chapter is concerned with progress in a current and very important aspect of space technology: communication relay satellite techniques. A pioneer in his field, the author demonstrates the need for long range communications, how the demand for communications can be satisfied, and how satellite communication systems operate. The major design factors in satellite systems are then reviewed, following which information is given on actual and proposed systems. Of particular interest is the fact that submarine cables can provide satisfactory service for most communication requirements. Consequently, to compete with the cable, a communications satellite system must satisfy four principal criteria which the author gives as its ability to provide (1) a bandwidth equal to or greater than that of the cable; (2) high quality, reliable service (high signal-to-noise ratio, small outage time); (3) the capability of being interconnected with major population centers on the Earth; and (4) a competitive cost structure.

FOREWORD

ix

A comprehensive review of solid propellant rocket technology follows, treating such subjects as factors affecting propellant performance, grain design, ignition, testing, propellant processing, inspection, quality control, grain defects, and thrust vector control and thrust termination. An interesting section is included on the use of solid rockets to power large space carrier vehicles. With the accelerated development of the 1,500,000lb liftoff thrust Saturn 1 and the 7,500,000-lb liftoff thrust Saturn 5, both liquid propelled carriers, the solid motor seems somewhat eclipsed. Nevertheless, it does hold considerable promise as an important first-stage propulsion system. Its success as a booster for the Titan 3 carrier will determine to a great extent its future. The advantages, disadvantages, and prospects of large solid propelled carriers are treated by the authors. A topic of commanding influence on manned space flight is studied by a researcher at the University of California Medical Center. In a fifty-odd page chapter on the environmental control of manned space vehicles such matters as atmospheric composition, atmosphere and life systems, inert gases, carbon dioxide, water, and toxic products are described. Attention is drawn to the fact that "besides atmosphere, food, and varying degrees of reduction and recycling of the products of metabolism, the space vehicle and its biological payload must achieve a balanced energy state compatible with life support in respect to heat storage and quantum exchange." The bioenergetics of the steady state is covered, the term "steady state" therein being defined as a state in which the energetic inputs and outputs of a system are equal. The author then proceeds to the problems of environmental control, examining both theory and applica­ tions. In the section devoted to this subject a basic framework of theory is provided "for specifying the characteristics of an environmental control compatible physiologically with requirements within the manned space vehicle." Turning to a completely different subject, three authors treat the impressive theme of navigation and geodetic satellites. Not only are terrestrial applications examined but many potential lunar and planetary relationships are conceived and explained. The chapter is divided into four principal parts, dealing, in turn, with (1) the navigation satellite (historical review, early concepts, and the Transit system); (2) the geodetic satellite (historical, specific examples of early satellites, methods of tracking, and the Anna system); (3) lunar and planetary navigation; and (4) the investigation of the gravitational fields of the Moon and planets with artificial satellites. Existing information on the major worlds of the Solar System is reviewed, following which anticipated characteristics, dynamics of orbiters, and data analysis and tracking requirements are covered. It is

X

FOREWORD

concluded that a "lunar orbiter should have readily observable perturba­ tions which will add significantly to our knowledge of the Moon's interior. . . . Satellites of Mercury, Venus, and Mars would also yield significant geophysical information as to their interior." It turns out, however, that satellites would be "considerably less sensitive measuring tools of the major planets. . . ." The final chapter in Volume 5 is by an eminent astronautical re­ searcher who has for years methodically inquired into his subject of orbital operations. He begins his chapter with the clear, declarative sentence: "Orbital operations are a prerequisite for any advanced space flight capability." He then shows that there are two classes of such operations, the first being those that must be undertaken because no alternative exists within the state of the art, and the second, those that are not dependent on the state of the art. In succeeding sections the author describes the role of orbital opera­ tions in astronautics, gives definitions of terms, and treats mission and systems integration in orbital operations, station keeping, intercept, delivery, pickup, rendezvous, fly-by, interorbital transfer, and typical orbital launch operations. The text is not limited to operations in Earthdominated space; operations in lunar, interplanetary, and planetary space also receive attention. FREDERICK I. ORDWAY,

Huntsvillej Alabama September 1963

III

Contents of Previous Volumes Volume 1 Interplanetary Rocket Trajectories DEREK F. LAWDEN

Interplanetary Communications J . R. PIERCE and C. C. CUTLER

Power Supplies for Orbital and Space Vehicles JOHN H . H U T H

Manned Space Cabin Systems EUGENE B . KONECCI

Radiation and Man in Space HERMANN J . SCHAEFER

Nutrition in Space Flight ROBERT G. TISCHER

Appendix. A Decimal Classification System for Astronautics HEINZ HERMANN KOELLE

Volume 2 Experimental Physics Using Space Vehicles CHARLES P. SONETT

Tracking Artificial Satellites and Space Vehicles KARL G. HENIZE

Materials in Space FREDERICK L . BAGBY

Plasma Propulsion Devices MORTON CAMAC

Electrostatic Propulsion Systems for Space Vehicles ERNST STUHLINGER and ROBERT N. SEITZ XV

Xvi

CONTENTS OF PREVIOUS VOLUMES

Attitude Control of Satellites and Space Vehicles ROBERT E. ROBERSON

Volume 3 The Role of Geology in Lunar Exploration JACK GREEN and JACK R. VAN LOPIK

Venus as an Astronautical Objective PATRICK MOORE and S. W. GREENWOOD

Mars as an Astronautical Objective SEYMOUR L. H E S S

The Exploration of Mercury, the Asteroids, the Major Planets and Their Satellite Systems, and Pluto RAY L. NEWBURN, JR.

Interplanetary Matter EDWARD MANRING

Structures of Carrier and Space Vehicles A. ALBERI and C. ROSENKRANZ

Advanced Nuclear and Solar Propulsion Systems WILLIAM C. COOLEY

Human Factors: Aspects of Weightlessness PAUL A. CAMPBELL

Volume 4 Doppler Effect of Artificial Satellites J . MASS and E. VASSY

On the Possibilities of the Existence of Extraterrestrial Intelligence ROGER A. MACGOWAN

The Development of Multiple Staging in Military and Space Carrier Vehicles H. E. NYLANDER and F. W. HOPPER

Spacecraft Entry and Landing in Planetary Atmospheres MAURICE TUCKER

Development of Manned Artificial Satellites and Space Stations SIEGFRIED J . GERATHEWOHL

On the Utilization of Radioactive Elements as Energy Sources for Spacecraft Propulsion J . J . BARRE

Supplementary Monograph

Supplement 1 : 0 . H . LANGE AND R. J . STEIN

Space Carrier Vehicles: Design, Development, and Testing of Launching Rockets, 1 9 6 3

xvii

Astronautical Investigations of the Sun R . GRANT ATHAY Department

of Astrophysics & Atmospheric University of Colorado and High Altitude Observatory Boulder, Colorado

Physics

I. Introduction II. Some Unsolved Solar Problems A. Thermodynamic Structure B. Geometrical Structure C. Velocity Fields D . Flares and Other Transient Phenomena E. Solar Magnetic Fields F. Corpuscular Emission III. The Role of Space Observations in Solar Physics A. General Considerations B. Ultraviolet and X-ray Radiation C. Corpuscular Radiation and Interplanetary Magnetic Fields . . . . IV. Demands on Laboratory Physics A. Spectroscopy B. Atomic Parameters References

1 3 3 6 7 9 9 10 12 12 13 15 17 17 18 18

I. Introduction

Astrophysical and geophysical research from extraterrestrial observa­ tories has opened a new era in the study of solar phenomena. Balloon, rocket, and satellite-borne instruments have already recorded many data that would otherwise be inaccessible. New phenomena have been discov­ ered and previously known ones have been observed with greater preci­ sion. This new avenue for research and its demonstrated successes have had a major impact on the entire scientific community interested in the Sun in any of its remarkably varied aspects. Yet, one has the feeling that we are only beginning; that we are still infants in our concept and under­ standing of solar phenomena. For every problem related to the Sun that we claim to understand, we 1

2

R. GRANT ATHAY

can name many related problems that at best are only poorly understood. New solar phenomena, recently discovered, emphasize that in all prob­ ability there is much happening on the Sun of which we are still totally unaware. Our attitude towards the space exploration of the Sun, then, is that it should be used as a means of pioneering into new realms of solar physics, both by searching out suspected phenomena that are shielded from ground-based observations by our terrestrial atmosphere and by seeking new phenomena as a means of broadening the scope of our re­ search efforts in both astrophysics and geophysics. Because of our lack of basic understanding of many solar phenomena, astronautical investiga­ tions of the Sun will continue for some time to serve as a guide in groundbased solar research almost as much as the ground-based research is able to guide astronautical research. Problems related to astronautical investigations of the Sun may be approached in alternative ways. In this chapter I shall attempt to outline the problems that, from our current vantage point, seem important irre­ spective of whether they are best solved through experiments carried aboard space vehicles. Following an outline of the problems, I shall at­ tempt to indicate those that may be attacked by experiments utilizing space vehicles. I am not well qualified to judge a particular space experi­ ment with regard to either its engineering or economic feasibility. Thus, in suggesting problems to be considered by those involved in astronautical research I do not necessarily imply that such problems are best solved in this way. However, in so far as I am able, I shall indicate the particular problems that seem most profitably solved through recourse to observa­ tions acquired beyond the confines of Earth's gaseous atmosphere and magnetic field. I t is beyond the scope of this chapter to define and describe all terms and concepts used in solar physics. The reader is referred to more exten­ sive discussions of the Sun for this purpose [1-6]. In the following discussion, the term "solar atmosphere" is used col­ lectively to include all visible layers of the Sun, including the photo­ sphere, chromosphere, and corona. When the Sun is observed at any particular wavelength all layers of the atmosphere contribute to the radiation. However, the great bulk of the observed radiation arises within a rather thin shell in the atmosphere, which we refer to as the "solar surface." Because the solar atmospheric opacity varies with wave­ length, the geometrical depth of the solar surface with respect to a fixed reference sphere varies with wavelength. Furthermore, since the Sun is a spherical object, the optical path from an observer to the Sun traverses the solar atmosphere in a direction radial to the Sun at the center of the solar disk and tangential to the Sun at the solar limb. Thus, a spherical

ASTR0NAUTICAL INVESTIGATIONS OF THE SUN

3

shell concentric with the Sun presents a greater opacity when viewed at the limb than when viewed at the center of the disk. The "solar surface" then increases in its radial distance from the center of the Sun as the optical path moves from center to limb on the solar disk. In the visual regions of the spectrum the brightness of the Sun decreases from centerto-limb, which is interpreted as an outward decrease of temperature.

II. Some Unsolved Solar Problems A. Thermodynamic Structure

Much of the current emphasis in solar physics is directed toward determining thermodynamic parameters characterizing a particular level in the "quiet" solar atmosphere or in a particular phenomenon of solar activity, e.g., flares, prominences, sunspots, etc. In a first approach to understanding the solar atmosphere, one assumes that the energy equi­ librium is established by radiative processes alone and that both hydro­ static and local thermodynamic equilibrium apply. Under such conditions, it is sufficient to specify temperature, density, and chemical composition. Furthermore, these parameters are relatively easy to ascertain from data obtained in the visual-photographic regions of the spectrum. If the Sun obeyed these assumptions, in fact, it would provide little of interest to astronautical research. It is well known that the solar atmosphere departs from the foregoing simplifying assumptions, but there is still much controversy over the magnitude, and in some cases even the direction, of such departures. As an illustration of the difficulties encountered in the three stated assump­ tions, we consider the following approximate conditions in the solar at­ mosphere. In the photosphere the opacity is due largely to the H"~ ion. The optical depth, T, to a given point in the solar atmosphere measured along the line from the observer to the point in question is r„ = avnijnekT L,

(1)

where av is the absorption coefficient, v is frequency, nH is the density of neutral hydrogen, nekT is the electron pressure and L is the effective at­ mospheric thickness. Tables of ap have been computed for the case of local thermodynamic equilibrium [7], which has been shown to be a good approximation for the establishment of the H ~ equilibrium [6, 8]. Within the photosphere hydrogen is predominantly neutral and most of the electrons are furnished by heavy atoms, whose abundances are 4 about 10~ that of hydrogen and which are predominantly singly ionized. 4 2 Thus, we may replace the product nHne by 10~ nn . The effective at-

4

R. GRANT ATHAY

mospheric thickness for the opacity at A5000 is about 60 km when the Sun is viewed radially and about 15,000 km when viewed tangentially. Using computed values of ap, and T = 6000°, we find T = 1 in the radial 17 -3 direction for nHzz2 X 10 c m and in the tangential direction at the 16 3 limb for nH ~ 10 cm~ . The difference in geometrical depth along the surface r = 1 is of the order of 300 km from the center of the disk to the limb. Therefore, if we use the limb as a reference point for measuring depth, the gas pressure has increased by about a factor of 20 at —300 km. The temperature distribution in the photosphere can be inferred from the measured brightness of the solar disk from center-to-limb, the socalled limb darkening. At - 3 0 0 km, we find T « 6000°, and at - 1 0 0 km we find T ~ 4700°. These differences in T will not change substantially the densities inferred above. Temperature and density distributions beyond the edge of the solar disk can be obtained, within certain limitations, from eclipse data. With­ out enumerating the difficulties of such an analysis we give approximate representative values in Table I [6]. TABLE

I.

APPROXIMATE REPRESENTATIVE VALUES OF TIH, ne, AND Pg

Te

(GAS PRESSURE) IN THE SOLAR ATMOSPHERE

Height (km) -300 UH ne Te Pg

3

17

500

0

-100 16

16

10 12 (cm""3 ) 2 X 10 13 3 X 10 12 (cm' ) 10 2 X 10 3 X 10 4300 3 (°K) 6000 4700 4 2 (dyne cm" ]1 2 X 10 6 2 X 10 6 X 10

14

1000

10,000 12

10 2 X 10 11 5 X 10" 3 X 10 6000 7000 80 2

9

10 9 10 6 1 X 10 0.1

100,000

—8

10 6 1 X 10 0.01

Several important conclusions can be drawn from Table I. Within the photosphere, h < 0, temperature, density, and pressure distributions are quite consistent with the assumptions of hydrostatic, radiative, and local thermodynamic equilibrium (the latter as applied to H~~). However, near the top of the photosphere the temperature of free electrons, jTe, begins to increase outwards, and between 1000 km and 10,000 km Te jumps ab­ ruptly and violently. The assumption of radiative equilibrium requires a monotonic outward decrease in Te. Hence, we must drop the assumption of radiative equilibrium as soon as Te departs significantly from its ex­ pected monotonic decrease, i.e., near h = 0. At heights > 500 km, some nonradiative source of energy must dominate the energy input in order to produce the large rises in Te.

ASTRONAUTICAL INVESTIGATIONS OF THE SUN

5

An additional conclusion embodied by the results in Table I, but which is not self-evident, is that at all heights above h = 0 the assumption of local thermodynamic equilibrium fails severely when applied to spec­ tral lines. For example, the chromosphere is opaque to several strong Fraunhofer absorption lines and the centers of these lines are formed at h > 500 km where Te > 6000°. Yet, these same lines radiate as a black body at a radiation temperature near 4000°. Similarly, the center of the Lyman-a emission line of hydrogen is formed in the region where Te ~ 70,000-90,000 [9, 10], but it has an equivalent black body radiation tem­ perature of about 7000°. Perhaps the most crucial unsolved problem in the solar atmosphere is the nature and origin of the nonradiative source of energy, which pro­ duces the outward increase in temperature and, indirectly, leads to the violent departures from thermodynamic equilibrium. There have been many speculations on the nature and source of this energy, the most plausible of which tie the energy source to the subphotospheric convec­ tion zone. Within this convection zone, wave disturbances must be gen­ erated and propagated outward. Energy carried by these waves is ex­ pected to traverse the photosphere with but slight absorption, but it is also expected to be heavily absorbed in the chromosphere and corona. The most complete model of such an energy supply is that given by Osterbrock [11]. Qualitatively, this type of theory seems adequate. Nevertheless, the final test must be on a quantitative basis; at present we can specify the energy requirements of the chromosphere and corona only to a very rough approximation. In order to specify the energy requirements of the outer solar atmosphere we must have a far better description of all the rele­ vant thermodynamic properties of the atmosphere than we are now able to extract from available data. Temperature, density and pressure dis­ tributions in the solar atmosphere are known in much greater detail than those given in Table I. About all we can say regarding energy require­ ments with any degree of certainty is that the chromosphere requires more nonradiative energy to maintain it than does the much hotter co­ rona, and that the total energy required is at least as large as the ultra­ violet and X-ray flux for A < 1800 A. The extreme ultraviolet and X-ray emission of the Sun, which has been the focus for much of the rocket and satellite experimentation, arises in the chromosphere and corona. Many of the lines in this part of the spectrum are expected to arise, in part, in the regions of rapid tempera­ ture rise separating these two regions of the atmosphere. Thus, detailed spectroscopic data for these lines are crucial to the proper understanding of the thermodynamic structure and the energy balance. We shall return

6

R. GRANT ATHAY

to this point, as well as others we have mentioned, in Sect. I l l of this chapter. Even though the photosphere is known to be close to thermodynamic equilibrium for the formation of its continuum spectrum, it does not nec­ essarily follow that the same conditions are valid for its line spectrum. In fact, Pecker [12] and his co-workers have shown with a reasonable degree of assurance that many of the faint Fraunhofer lines depart from thermodynamic equilibrium. The main questions at stake with regard to the photospheric problems related to thermodynamic equilibrium are the precise specifications of the thermodynamic properties of the upper photosphere and the derivation of chemical abundances of elements. The latter could have far-reaching effects in astrophysics if it is conclusively demonstrated that chemical abundances determined from Fraunhofer line data (by assuming thermodynamic equilibrium) are shown to be signifi­ cantly in error, as has already been shown for the Fraunhofer lines formed in the chromosphere. B. Geometrical Structure

It is implicit in the assumptions of radiative and hydrostatic equilib­ rium that the solar atmosphere is spherically symmetric. All available data sharply contradict this conclusion. In the photosphere, granulation, faculae, sunspots, and the so-called "wiggly lines" in the Fraunhofer spectrum manifest widespread departures from spherical symmetry. The lower chromosphere is characterized by a "mottled" appearance and pro­ nounced wiggly-line structure when observed in the strong Fraunhofer lines. The upper chromosphere is in a continual state of upheaval caused by spicule eruptions. In fact, at heights above 3000-4000 km the chromo­ sphere (as we now know it) consists solely of the geyser-like spicules that appear at a given location about once a day and have lifetimes of only a few minutes. Streamers, loops, plumes, and condensations typify coronal structure and give vivid illustrations of departures from spherical symmetry. Dense clouds of prominence material imbedded in the corona represent further extreme departures from spherical symmetry. Without doubt one of the most pressing problems of solar physics is an accurate specification of the geometrical structure throughout the entire atmosphere. All interpretations of the radiant energy of the solar atmosphere, whether they be in terms of thermodynamic variables, ve­ locity fields, or chemical abundances of elements, depend upon some understanding of the atmospheric geometry. Photospheric geometry, and in some instances chromospheric geome­ try, can be studied in projection against the solar disk. In the chromo-

ASTRONAUTICAL INVESTIGATIONS OF THE SUN

7

sphere we have the added advantage of being able to study the geometry as seen at the solar limb. For the corona and for much of the chromo­ sphere we have essentially only this latter type of observation, and the problem of inferring the geometry is akin to that of studying the trees of a forest by standing far away and viewing the forest from the side. The ultraviolet, X-ray, and radio spectrum provide a unique possi­ bility for studying chromospheric and coronal geometry as seen "looking down from the top" where individual features should stand out more clearly. Unfortunately, these regions of the spectrum present serious problems of their own in getting the necessary angular resolution to see the fine details of solar structure. Nevertheless, the possibility exists in these regions of the spectrum for getting this basic information, and there seems little hope of getting it accurately otherwise. Our thermodynamic picture of the solar atmosphere is intimately associated with the geometrical structure assumed. Many models, based on specific departures from spherical symmetry, have been proposed. It is perhaps safe to say that these models indicate the directions of the effects they produce on the thermodynamic variables, but it would be hardly fair to say that they accurately indicate the magnitude of these effects. At present, it appears that the "average" conditions in the photo­ sphere and chromosphere below 500 km do not depart strongly from hydrostatic equilibrium. At all heights in the corona and in the chro­ mosphere above about 500 km, the conditions of hydrostatic equilib­ rium are strongly violated, and this concept is of little aid in constructing models. When we add to the complexity of departures from spherical symmetry by abandoning the concepts of thermodynamic equilibrium and by allowing unknown energy transport mechanisms, the problems of inferring the thermodynamic properties of the atmosphere are enor­ mously compounded. The solution to these problems will demand accurate knowledge of the complete solar spectrum and its variation from point to point in the solar atmosphere. C. Velocity Fields

The recent discovery of vertical oscillatory motions in the photosphere and lower chromosphere [13, 14] represents a major advance in solar physics. The proper interpretation of these motions, however, is still far from being clear. A line in the solar spectrum may exhibit a Doppler shift either by a coherent motion of the entire atmosphere in the regions where the line is formed or by movement of a small vertical segment of the atmosphere. The magnitude of the resultant shift in the line for a

8

R. GRANT ATHAY

given velocity may differ markedly in the two cases; and, consequently, an observed Doppler shift may allow a wide range of interpretation. There are undoubtedly cases where the interpretation is straightforward and simple. However, in most cases the interpretation of observed velocity fields in the solar atmosphere demands a reliable description of the geometrical and thermodynamic properties of the atmosphere. A possible example of the difficulty in interpreting solar velocity fields is evidenced by the spicule phenomenon. Spicules apparently originate near the 1000-km level in the chromosphere [6] and propagate upward with velocities of about 25 km/sec when observed at the limb. They move predominantly radially to the Sun, but some are inclined enough to the radial direction to produce a Doppler shift in their spectral lines observed at the limb. By this we know that the motions are real mass motions and not "excitation waves" of some sort. These same spicules, when observed against the center of the solar disk, should shift their absorption profile in the Balmer-a line of hydrogen by about 0.55 A, and they are numerous enough that they should be relatively easily observed. Many observers have claimed to see disk features observed in the Balmer-a line of hydrogen identifiable as spicules, but no one, to my knowledge, has reported Doppler shifts approaching those expected for spicules. The difficulty is very likely associated with the fact that the opacity of a spicule in the Balmer-a line is of the order of ten, whereas the regions through which the core of the Fraunhofer line is formed have 4 a total opacity of about 10 . The effect of a spicule on the undisturbed line profile will be further complicated by the fact that the absorption profile of the spicule in the Balmer-a line is much broader than the absorption profile in the regions where the line is normally formed and by the additional fact that the source function may be much different also. The importance of velocity fields in the over-all problems of solar physics should not be minimized. The solar wind, even if only a breeze, is of major importance in the physics of the interplanetary medium as well as in the corona. Matter transported upwards in spicules is enough 7 to replace the entire corona in a few hours time and much more than is required by the solar wind. The kinetic energy of their motion is com­ parable to the energy radiated by the corona. By comparison, the granu­ 5 3 lations carry about 10 more mass flux and about 10 more energy flux than spicules. The solar granulation is a manifestation of convection cells in the convectively unstable layers below the photosphere, which in all proba­ bility is the origin of the nonradiative energy dumped into the chromo­ sphere and corona. Are the spicules directly related to granules, or are

ASTRONAUTICAL INVESTIGATIONS OF THE SUN

9

they secondarily related as the wake of Shockwaves produced as subsonic waves in the convection zone and accelerated outward as suggested by Osterbrock [11]? Or, are they a completely independent phenomenon produced by some dynamic instability in the low chromosphere? Once again, the answer to these questions hinges on a better understanding of the thermodynamics, geometry, and energetics of the solar atmosphere, and, hence, on the acquisition of sufficiently detailed data from extra­ terrestrial observatories. D. Flares and Other Transient Phenomena

All that has been said about the difficulties of understanding the quiet Sun becomes even more pronounced in the case of solar flares and most other transient solar phenomena. The energetics, geometrical and thermodynamic problems of the so-called quiet Sun are still present in flares. In addition, two added complexities appear: flares are a timedependent phenomena and they either produce or are related to the production of copious high-energy particles. Magnetic fields are closely associated with and are involved in all types of solar activity, but we may not be correct in believing that they are not similarly involved in phe­ nomena of the quiet Sun. Again, one of the great difficulties in any attempt to understand flares and related phenomena is a basic lack of the proper kinds of data. Much emphasis has been placed on detecting and classifying flares because of their important geophysical consequences and because of their marked effects on the interplanetary medium. Relatively little has been done in a serious attempt to understand the flare phenomenon. The flare phenomenon manifests itself in the electromagnetic spectrum from radio frequencies to X-rays and in the particle spectrum from a few hundred ev into the Bev range. In the simple world of thermodynamic equilibrium all of these effects could be understood with limited observa­ tional data. In the actual case they cannot. We must know the complete characteristics of many selected features of the spectrum, a problem that presents one of the greatest challenges to astronautical investigations of the Sun. E. Solar Magnetic Fields

Our knowledge of solar magnetic fields is severely limited by the fact that we have thus far observed in some detail only the photospheric fields, and then only the line-of-sight components. The presence and sometimes shape of chromospheric and coronal magnetic fields are made manifest

10

R. GRANT ATHAY

by polarized radio signals, coronal features, prominence trajectories, and alignment of both chromospheric spicules and the mottling observed in strong Fraunhofer lines. Leighton [15] and Howard [16] have discovered relationships between photospheric magnetic fields and chromospheric plages, and more recently Severny and Zirin [17] have reported observa­ tions of magnetic fields in prominences. All of the above-mentioned information on solar magnetic fields refers to regions relatively near the solar surface and, even then, the information is very incomplete. For many purposes we need to know magnetic field strengths and configurations in the neighborhood of the Sun, but somewhat farther out than present data permit. Just how we will get such information is not clear, but observations from space vehicles of the polarized component of the white light corona and of the location and polarization of enhanced emission regions of the Sun as seen at wavelengths in the radio spectrum too long to penetrate the ionosphere would be of valuable assistance. Similarly, solar probes within the orbit of Mercury for measuring magnetic fields and plasma fluxes would be of great value. There have been several theoretical, semi-theoretical, and semi-em­ pirical models proposed for the configuration of the solar magnetic field in interplanetary space. Solar cosmic ray data offer valuable clues to the nature of the fields, but they have not yet led to a unique, unambiguous model. Direct measurements of solar plasma arriving in the vicinity of Earth, however, will continue to provide much of the basic information on interplanetary fields. F. Corpuscular Emission

Sporadic ejection of clouds of energetic particles from the Sun, largely in association with major solar flares, has been known for several decades. However, only in the last few years has there been any measure of detailed knowledge of the number and energy spectrum of these particles. Even now our knowledge is still confined to essentially pre­ liminary stages of formulation of ideas based on fragmentary data. As observational sophistication has increased, so has the number and character of known solar corpuscular events. In general, the sporadic events that produce solar cosmic rays, polar cap proton events, low latitude aurorae, and sudden commencement magnetic storms tend to parallel the occurrence of major flares, with the exception that the cosmic ray events seem to become relatively rarer near the maximum of solar activity [18]. This avoidance of solar cosmic ray events to high solar

ASTRONAUTICAL INVESTIGATIONS OF THE SUN

11

activity has suggested to some that it results from a great disorder in the interplanetary magnetic fields that prohibits these particles from reaching the Earth. Additional evidence supporting this conclusion comes from the general decrease of galactic cosmic rays at sunspot maximum. On the other hand, Takakura and Ono [19] have suggested that intense radio outbursts at centimeter wavelengths, which are correlated with the cosmic ray events, show the same tendency to avoid sunspot maximum, and, therefore, that the effect may be mainly of direct solar origin. The only known source of energy that seems plausible as an accelera­ tion mechanism for energetic solar particles lies in solar magnetic fields. The close association of the particle events with flares implies a close association with complex magnetic field configurations. Recently ob­ served high-velocity disturbances propagating outward from some flares during an "explosive" phase [20, 21], and the association of this explosive phase with 10-cm radio bursts [22], points to a possible closer linkage between optical and particle events than had been known previously. Fast drift and type IV radio events are believed to be indicative of high energy particles in the solar atmosphere; and, in fact, type IV radio events correlate closely with polar proton events. These data provide valuable means for detecting energetic solar events and of studying their characteristics. However, they do not replace the need for direct obser­ vations of the solar plasma in the vicinity of Earth and as near the Sun as possible. The quiet as well as the active Sun emits corpuscular radiation. Al­ though these particles are far less energetic than the active Sun particles, their importance in solar and interplanetary problems cannot be over­ looked. The computation of the solar wind [23] depends critically upon the assumed thermodynamic properties of the corona in external regions where we have only scanty observational data. Thus, there is little that the solar physicist can contribute to the understanding of this phe­ nomenon without a more complete understanding of the thermodynamics and energetics of the outer corona. Again, our best hope of getting this information is from space observatories. Direct observation of the corpuscular radiation of the quiet and active Sun from space observatories has long been the dream of the solar physicist. The mysterious " M " regions of the Sun that produce recurrent geomagnetic storms during periods of low solar activity are still unexplained, and they have not been convincingly identified with any known solar phenomenon. Similarly, the nature of coronal streamers, which represent extensions of coronal plasma far into interplanetary space, is not at all understood. These problems, together with spicule

12

R. GRANT ATHAY

eruptions and surge- and spray-type prominence ejections, may relate closely to the problems of the solar wind and to the sporadic accelerations of energetic particles. None can be claimed to be adequately understood. III. The Role of Space Observations in Solar Physics A. General Considerations

The foregoing outline of some of the unsolved solar problems is 1 necessarily brief and sketchy. However, I intend that it convey to the reader only the nature of the problems and some feeling for their scope. In most practical cases, specific space experiments should be designed so that they answer specific needs and questions. The experimenter must be careful that his experiment is capable of answering questions that need answering and that cannot be more easily and more economically an­ swered. Thus, if the foregoing outline either encourages interest in solar problems or aids in choosing the types of problems to be investigated it has served its purpose. Some of the problems we have discussed can be answered by suitable observations made only once. Others require continuous monitoring. Nearly all imply some recourse to observations made outside the E a r t h s atmosphere. There are three basic reasons to observe the Sun from outside the atmosphere rather than from ground-based observatories. (By outside the Earth's atmosphere we mean sufficiently high to escape the particular atmospheric effect that prohibits ground-based observations.) These rea­ sons are: (1) to observe solar radiations that do not penetrate to the ground; (2) to eliminate the bright sky caused by scattering of sunlight in our relatively dense atmosphere; and (3) to escape seeing effects caused by dense atmospheric air currents of different index of refraction. The first two of these are the most important for the majority of problems we have mentioned. Except in rare instances the entire spectrum of corpuscular radiation from the Sun is unobservable by direct measurement from the ground. At wavelengths shorter than about 3000 A, the Earth's atmosphere becomes opaque and the incoming solar radiation is absorbed in the ozone and ionospheric layers. At the opposite end of the spectrum the ionosphere becomes opaque to radio waves longer than about 75 meters at sunspot minimum and at somewhat shorter wavelengths at sunspot maximum. Even though the atmosphere is effectively transparent in the 1

F o r a more complete discussion of these problems and their interrelationships, the reader is referred to standard texts on the Sun and to a review article on solar indices by Athay and Warwick [ 2 4 ] .

ASTRONAUTICAL INVESTIGATIONS OF THE SUN

13

optical regions of the spectrum, there are many solar phenomena not yet successfully observed because of a low signal-to-noise ratio, caused by our bright terrestrial sky and an intrinsically faint phenomenon on the Sun. Many of these phenomena can be observed at the time of total solar eclipse. However, eclipses, and particularly those with promise of clear skies and favorable geographic location, are relatively rare; and it is impossible to build up a satisfactory picture of the structure and evolu­ tion of these faint solar features. Examples of this type of phenomena are the corona (beyond about 1.5 solar radii), the outer extensions of prominence material ejected from the Sun and, hopefully, solar plasma clouds of the type that produce geomagnetic storms. Atmospheric seeing effects are particularly troublesome in attempts to measure relative brightness of solar features that are either sharply bounded in space or that have steep gradients in brightness. Both of these problems are present in attempts to study sunspot structure and chromospheric features at the solar limb. Assuming that we have the necessary engineering skill, we could greatly improve the quality of this type of solar data by observing from space vehicles, as has already been demon­ strated by Schwarzschild [25]. B. Ultraviolet and X-ray Radiation

Much of the groundwork in observing the ultraviolet and X-ray spectrum of the Sun has already been laid, but far more remains to be done. A logical sequence of steps for a systematic study of this part of the spectrum would be: (1) accomplish a scan of the spectrum with sufficient sensitivity and resolution to establish the continuum and indi­ vidual line intensities at all wavelengths; (2) select certain spectral lines for detailed studies of their profiles as dictated by their importance in helping to understand the thermodynamic and geometrical structure of the Sun, such as the resonance series lines of the ions of abundant ele­ ments; (3) monitor the spectrum in order to identify features that vary with solar activity; and (4) isolate these latter features and study their intensities and profiles as functions of time and spatial distributions on the solar disk. In practice, of course, we must modify this type of sequence to be consistent with limitations imposed by our skills in instrumentation and our opportunities for observing. However, one should not lose sight of the over-all programs because of these limitations. The Sun should be patrolled just for the purpose of identifying activity, as it has been in the past. But, too often, such patrol programs choke off research by absorbing a disproportionate share of the total effort. Space observations of the Sun will undoubtedly uncover many important aspects of solar

14

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activity that will beg for more routine patrolling than that in which we are now engaged. However, logic dictates that our primary effort should be to understand solar phenomena rather than merely watch them happen repeatedly. If we are to understand solar phenomena as we should, we must have detailed information on intensities and profiles of sufficient spectral features to specify conclusively all of the relevant parameters of the phenomena. The chromosphere-corona transition region, the upper chromosphere, the corona, and essentially all forms of solar activity can hardly be understood without detailed information on a number of spec­ tral lines and continua in ultraviolet and X-ray regions of the spectrum. 11 1 11 Lines, such as the first few resonance lines of hydrogen, Mg , He , He , IV VI C , O , and the stronger resonance lines of coronal ions, should be care­ fully observed with enough spectral resolution to give details of line profiles as soon as it is practical. The spectral resolution required for most of the lines of interest in order to obtain the necessary profile 5 information is of the order of A A / A < 5 x 1 0 ~ . Pioneer studies of the solar ultraviolet and X-ray spectrum, such as those carried out by Rense [26], Tousey [27], Friedman [28] and Hinteregger [ 2 9 ] , have already revealed some of the interesting features of the solar spectrum in this wavelength range. The spectrum has been mapped at moderate-to-low resolution down to wavelengths of about 60 A with absolute intensity measurements of all prominent features. However, many of the weaker lines are still unresolved from neighboring lines and accurate profile information is available only for the hydrogen Lyman-a line. For A > 250 A, most of the strong lines are identified, but the weaker lines observed are still mostly unidentified. For A < 250 A, essentially none of the observed lines has been identified [30]. At wave­ lengths below 60 A only broad band resolution has been achieved, and the detailed character of the spectrum is still relatively unknown. Many lines in the extreme ultraviolet and X-ray regions of the solar spectrum are expected to vary considerably with solar activity. A pre­ liminary report of variations in association with flares has been made from data obtained by the Orbiting Solar Observatory for the lines A504 11 xv of He , A284 of F e and A171 (unidentified) [ 3 1 ] . Variations for the integrated radiation from the solar disk are of the order of 10-30 per cent. A conservative estimate for the fraction of the solar disk showing en­ hanced radiation is 0.01 as an upper limit. Thus, we infer that in local regions near the flare the disk brightness has increased at least 10-30 times over its normal value and quite possibly by an additional factor of 10. In the X-ray region below 60 A, much larger percentage variations

ASTRONAUTICAL INVESTIGATIONS OF THE SUN

15

are observed, both with and without accompanying flares. For example, Chubb et al. [32] report that the X-ray flux at times of no flares increased from sunspot minimum to sunspot maximum in the 44-60 A band by a 3 factor of seven and in the 2-8 A band by a factor of at least 10 . X-ray flux changes associated with solar activity are characterized by a hard­ ened spectrum as well as increased flux; and, to date, measurable fluxes have been observed below 0.5 A only during flares. X-ray flux increases in excess of a factor of 20 have been observed in the 2-8 A band in the SR-1 (Greb 1) satellite in association with flares, and changes of similar magnitude have also been observed in association with surges and active region prominences [33]. The broad self-reversed cores of the chromospheric hydrogen Lyman-a n line and the M g resonance lines are of great interest in solar physics. Detailed shapes of other lines originating in the chromosphere and chromosphere-corona transition, such as the hydrogen Lyman-/? line, the 11 VI H e Lyman-a line at A304 and the O resonance lines, would be of equal interest if data were available. Direct photographs of the solar disk in the Lyman-a [34] and in the X-ray region of the spectrum [35] showing localized areas of enhanced emission over active regions are exciting and commendable steps in the right direction. Much more work of this type is needed, however, espe­ cially if these space observations can be coordinated with ground-based observations in different regions of the spectrum. Lines of the solar spectrum particularly sensitive to solar activity, such as those recently reported [31], need careful studies of intensities and profiles both during periods of variation and periods of relative quiet. Many of the solar observations associated with the ultraviolet and X-ray spectrum need exploratory programs that can be carried out on rocket sounding vehicles. However, observations requiring either long integration time, accurate time resolution, or long time coverage must rely on satellite vehicles. C. Corpuscular Radiation and Interplanetary Magnetic Fields

Direct observations of solar corpuscular radiation must be made outside the terrestrial atmosphere. Indirect studies of these phenomena via such means as geomagnetic, auroral, and ionospheric variations leave many questions unanswered, as was brought dramatically to light by the discovery of the Van Allen radiation belts. No amount of indirect study will satisfactorily replace direct observations, at least until after enough direct observations have been made to lead to a clear understanding of the complicated interrelationships between geophysical and solar events.

16

R. GRANT ATHAY

Solar plasma in transit from Sun to Earth must be intimately asso­ ciated with magnetic fields, regardless of whether such fields are carried along by the plasma or semi-permanently lodged in interplanetary space. Thus, corpuscular radiation and interplanetary magnetic field phenomena should be simultaneously observed whenever possible. Satellites with eccentric polar orbits and long range solar probes are particularly useful for such studies. Orientations and intensities of magnetic fields and the direction of arrival and energy spectrum of particles (both far from Earth and as they interact with the geomagnetic field) are necessary to adequately understand the nature of solar plasma clouds and their asso­ ciated terrestrial effects. Solar corpuscular radiation has quiet and active components, just as the ultraviolet and X-radiation. However, corpuscular radiation is far more variable and unpredictable than the short wavelength electromag­ netic spectrum. So little is known about the corpuscular spectrum at this time that essentially any carefully done experiment to detect plasma clouds and magnetic fields in interplanetary space is of great value. The quiet Sun, or so-called "solar wind," component of corpuscular radiation has been determined from Lunik 2 [36] and Explorer 10 [37] 8 -2 -1 observations to have a proton flux of about 4 X 10 c m s e c with energies of about 500 ev, which corresponds to proton densities of about - 3 20 c m . The direction of flight of the protons at the positions of the space vehicles was not precisely determined, but was shown to be from approximately the direction of the Sun. Direct plasma from the Sun was first observed by Explorer 10 at 21.5 Earth radii. Beyond this distance the flux varied periodically from about the value quoted above to the 6 -2 - 1 minimum detectable level of 5 X 10 c m sec " , with a wavelength of 6 roughly 2 X 10 km. Magnetic field measurements made by Explorer 10 [38] showed a similar variation in magnetic field intensity and direction, but exactly out of phase with the plasma flux. During periods of high plasma flux, the magnetic field was weak and variable and during periods of low plasma flux the field was relatively strong and oriented about 40 deg west of the Sun. The above-mentioned measurements of the solar wind were made at a moderately high level of the solar cycle. Long period variations through the sunspot cycle seem probable and should be looked for in future experiments. Measurements of sporadic clouds of solar plasma from space vehicles are still so rare as to preclude any meaningful summary. Expected energies of such clouds range from a few hundred electron volts into the Bev range and particle trajectories from monodirectional to isotropic. Maxi­ mum particle fluxes are still hard to estimate because of the paucity of

ASTRONAUTICAL INVESTIGATIONS OF THE SUN 11

-2

17

1

data, but fluxes of 10 c m sec""" are not out of the realm of possibility. At a distance of about 37 Earth radii Explorer 10 detected a sharp in­ 2 crease in proton flux of at least a factor of 10 over the background solar wind flux, and a hardening of the energy spectrum about 9 min after a sudden commencement, followed by a relatively weak geomagnetic storm, was observed at the surface of the Earth [37]. The magnetic field de­ tected by Explorer 10, however, began a slow increase at the time of the sudden commencement and then increased more rapidly with the arrival of the enhanced plasma flux [38]. During periods of high solar activity, sporadic plasma emission from the Sun is largely flare-associated. However, because of the wide range of energies and complexity of individual particle trajectories the plasma may arrive in the vicinity of Earth from immediately after the flare to as much as 3-4 days later. The solar wind, if present, will likely dominate the particle spectrum of times of low solar activity. M-region particles should appear period­ ically with 27-day recurrences during the waning part of the sunspot cycle. These plasma ejections, if truly of solar origin and not repre­ sentative of some sort of triggering of the outer Van Allen belts, should appear as more or less continuous streams rather than discrete clouds and should carry relatively large particle fluxes. Plasma clouds responsible for sudden commencement geomagnetic storms are of rather low energy per particle, but may contain relatively large total energies because of the number of particles present. The possi­ bility of observing such plasma clouds by optical means from satellite vehicles should not be overlooked. At an angular distance of 10 deg from 4 the Sun the zodiacal light is about 10~ as bright as the corona at 0.5 radii beyond the limb, and the flux of solar radiation is reduced by about - 2 1 0 . At the solar equator, 0.5 radii beyond the limb, the product of ne 17 - 2 and the effective path length, L, is about 10 c m . Hence, a plasma 15 -2 cloud 10 deg from the Sun in which neL ~ 10 c m should be as easily observed as the zodiacal light. For a cloud whose diameter equals the 4 solar radius, this requires ne ~ 10 , which does not seem improbably high. IV. Demands on Laboratory Physics A. Spectroscopy

One of the primary difficulties of analysis in the solar ultraviolet and X-ray spectrum is the proper identification of spectral lines. Laboratory wavelengths, intensities, and identifications of the vast majority of the lines observed are not available, Theoretical identification often require^

18

R. GRANT ATHAY

large, uncertain extrapolations from the last well-calculated or observed energy level separation. Techniques of spectroscopy and of image formation in this part of the spectrum are in urgent need of improvement from their currently crude state. Existing photometric techniques are also relatively crude and in need of improvement. B. Atomic Parameters

The entire solar line spectrum in the ultraviolet and X-ray regions is formed under conditions departing strongly from thermodynamic equilib­ rium. As a result, detailed rate processes in atoms must be known before intensities and profiles of lines can be reliably interpreted. Appropriate atomic parameters for determining these rate processes are known in only a few cases; and, even then, for only a few transitions in the most intensively studied atoms. Oscillator strengths (/-values) and collision cross-sections are required for all atoms and ions whose spectral lines are present. I t is not sufficient to know just those atomic parameters related to the immediate levels from which the spectral lines arise. Accurate rate processes must be known for all transitions important in establishing the equilibrium con­ figurations of the atom. These levels, in many cases, can be determined from preliminary calculations using only approximate rates. Again, how­ ever, such calculations are relatively rare and need to be extended.

References 1. Ellison, M. A. (1955). "The Sun and Its Influences." Routledge & Paul, London. 2. de Jager, D . (1959). "Handbuch der Astrophysik" (S. FlUgge, ed.), p. 112. Springer, Berlin. 3. Kuiper, G. P., ed. (1953). "The Sun." Univ. of Chicago Press, Chicago, Illinois. 4. Menzel, D . H. (1959). "Our Sun," rev. ed. Harvard Univ. Press, Cambridge, Massachusetts. 5. Shklovsky, I. S. (1962). "Fizika Solnechnoi Korony." Government Printing Office (Math, and Phys. Lit.), Moscow. 6. Thomas, R. N., and Athay, R. G. (1961). "Physics of the Chromosphere." Wiley (Interscience), New York. 7. Chandrasekhar, S., and Breen, F. H. (1946). On the continuous absorption co­ efficient of the negative hydrogen ion III. Astrophys. J. 104, 430-445. 8. Pagel, B. E. J. (1957). The emission of continuous radiation in stellar atmosphere. Astrophys. J. 125, 298-300. 9. Morton, D . C , and Widing, K. G. (1961). The solar Lyman-a emission line. Astrophys. J. 133, 596-605.

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10. Athay, R. G. (1960). The equilibria and ultraviolet spectra of H, H e l and H e l l in the solar atmosphere. Astrophys. J. 131, 705-716. 11. Osterbrock, D. E. (1961). The heating of the solar chromosphere, plages and corona by magnetohydrodynamic waves. Astrophys. J. 134, 347-388. 12. Pecker, J.-C. Series of papers beginning (1959). Ecarts A L'Equilibre et Abondances dans les Photospheres solaire et stellaires. I. Le Spectre du Titane Neutre —Ecarts a 1'E.T.L. Ann. astrophys. 22, 499-522. 13. Leighton, R. B., Noyes, R. W., and Simon, G. W. (1962). Velocity fields in the solar atmosphere. I. Preliminary report. Astrophys. J. 135, 474-499. 14. Evans, J. W., and Michard, R. Series of papers beginning (1962). Observational study of macroscopic inhomogeneities in the solar atmosphere. I. Velocity dis­ placements of Fraunhofer lines as a function of line strength and position on disk. Astrophys. J. 135, 812-821. 15. Leighton, R. (1959). Observations of solar magnetic fields in plage regions. As­ trophys. J. 130, 366-380. 16. Howard, R. (1959). Observations of solar magnetic fields. Astrophys. J. 130, 193201. 17. Severny, A., and Zirin, H. (1961). Measurement of magnetic fields in solar promi­ nences. Observatory 81, 155-156. 18. Warwick, C. S. (1962). Propagation of solar particles and the interplanetary magnetic field. J. Geophys. Research 67, 1333-1346. 19. Takakura, T., and Ono, N. (1962). Yearly variation in activities of outbursts at microwaves and flares during a solar cycle with special reference to unusual cosmic-ray increases. J. Phys. Soc. Japan 17, 207-210. 20. Moreton, G. C. (1959). Paper presented to Cleveland Meetings of Am. Astron. Soc. 21. Athay, R. G., and Moreton, G. C. (1961). Impulsive phenomena of the solar atmosphere. I. Some optical events associated with flares showing explosive phase. Astrophys. J. 133, 935-945. 22. Covington, A. E., and Harvey, G. A. (1961). Coincidence of the explosive phase of solar flares with 10.7 cm solar noise bursts. Nature 192, 152-153. 23. Parker, E. N. (1959). Extension of the solar corona into interplanetary space. J. Geophys. Research 64, 1675-1681. 24. Athay, R. G., and Warwick, C. S. (1961). Indices of solar activity. Advances in Geophys. 8, 1-83. 25. Bahng, J., and Schwarzschild, M. (1961). The temperature fluctuations in the solar granulation. Astrophys. J. 134, 337-342. 26. cf. Violett, T., and Rense, W. A. (1959). Solar emission lines in the extreme ultraviolet. Astrophys. J. 130, 954-960. 27. cf. Tousey, R. (1953). In "The Sun" (G. P. Kuiper, ed.), pp. 659-670. Univ. Chicago Press, Chicago, Illinois. 28. cf. Friedman, H., Lichtman, S. W., and Byram, E. T. (1951). Photon counter measurements of solar x-rays and extreme ultraviolet light. Phys. Rev. 83, 1025-1030. 29. cf. Hinteregger, H. E. (1961). Preliminary data on solar extreme ultraviolet radi­ ation in the upper atmosphere. J. Geophys. Research 66, 2367-2380. 30. Zirin, H., Hall, L. A., and Hinteregger, H. E. (1962). Analysis of the solar emission spectrum from 1300 to 250 A as observed in August 1961. Proc. 3rd Intern. Space Symposium, Washington, D. C. in press.

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31. Bearing, W. E., Neupert, W. M., and Lindsay, J. C. (1962). Preliminary solar flare observations with a soft x-ray spectrometer on the orbiting solar observa­ tory. Report X-614-62-29, Goddard Space Flight Center, Greenbelt, Maryland. Paper presented at COSPAR, May 1962. 32. Chubb, T. A., Friedman, H. A., and Kreplin, R. W. (1961). X-ray emission accompanying solar flares. Les Spectres des Astres dans L'ultraviolet Lointain, Mem. Soc. Roy. Sci. Liege [5] 4 (fasc. unique), 216-227. 33. Kreplin, R. W., Chubb, T. A., and Friedman, H. A. (1962). X-ray and Lyman-a emission from the Sun as measured from the N R L SR-1 satellite. Report of U. S. Naval Research Laboratory, Washington 25, D . C , 16 March. 34. Tousey, R. (1961). Photography of the Sun in Lyman-a and other wavelengths. Les Spectres des Astres dans L'ultraviolet Lointain, Mem. Soc. Roy. Sci. Liege [5] 4 (fasc. unique), 274-282. 35. Chubb, T. A., Friedman, H., Kreplin, R. W., Blake, R. L., and Unzicker, A. E. (1961). X-ray solar disk photograph. Les Spectres des Astres dans L'ultraviolet Lointain, Mem. Soc. Roy. Sci. Liege [5] 4 (fasc. unique), 228-240. 36. Gringauz, K. L, Bezrukikh, V. V., Ozevov, V. D., and Rybchinskii, R. E. (1960). "Iskusstvennye sputniki Zemli N o . 6." Izd. Akad. Nauk S.S.S.R., Moscow, 1961; see also: A study of the interplanetary ionized gas, high energy electrons, and corpuscular radiation from the Sun by means of the three-electrode trap for charged particles on second Soviet Cosmic Rocket. Doklady Akad. Nauk S.S.S.R. 131, 1301-1304. 37. Bridge, H. S., Dilworth, C , Lazarus, A. J., Lyon, E. F., Rossi, B., and Scherb, F. (1961). Direct observations of the interplanetary plasma, Proc. Intern. Con], on Cosmic Rays and the Earth Storm, Kyoto in press. 38. Heppner, J. P., Ness, N . F., Skillman, T. L., and Scearce, C. S. (1961). Magnetic field measurement with the Explorer X satellite. Proc. Intern. Conf. on Cosmic Rays and the Earth Storm, Kyoto in press.

Advances in Communication Relay Satellite Techniques R . P . HAVILAND Missile and Space Division, Philadelphia,

General Electric Pennsylvania

Company,

I. Introduction II. The Need for Long Range Communications A. Types of Communication Demand B. Factors Affecting Demand C. Estimation of Future Demand III. Satisfaction of Communication Demand IV. Basic Satellite Communication System Techniques V. Design Factors in Satellite Systems A. Coverage B. Time Delay C. Frequency Allocations and Sharing D. Attitude Control and Station Keeping E. Power Supply VI. Specific System Proposals VII. Status of Experimental Communication Satellite Programs A. Completed Programs B. Programs in Progress References

I. Introduction

21 22 22 23 23 26 28 31 31 33 36 37 40 41 44 44 44 45

[1-4]

The communication satellite field may be divided into sub-fields cover­ ing satellites used for different communication purposes. Three major types of these satellites distinguished so far are: (1) relay satellites, (2) broadcast satellites, and (3) distribution satellites. Relay satellites are intended to handle bulk message traffic between two points on the Earth's surface. The technique corresponds exactly to the communication service known as the point-to-point service. The term "relay" is used to indicate that the satellite receives messages from one point and relays it to the second point. Broadcast satellites are intended to provide program service to a large number of listeners scattered over a wide area. The service corre21

22

R. P. HAVILAND

sponds exactly to the present broadcast service, in which radio and TV are used to transmit programs to home owners. The distribution satellite is a recently recognized separate version which is intended to transmit specialized material from its source to a number of distribution points which "retail" the signal to the ultimate user. This service corresponds roughly to the operation of a network used in the distribution of TV programs, but is not limited to program material. For example, another type of data which requires distribution are weather forecasts. These are currently distributed by teletype to radio stations, airports, and newspapers, which then transmit the material to the end user. At the present time the major effort in communication satellites re­ lates to the relay or point-to-point service. The remainder of this chapter is devoted entirely to this service and will discuss the reasons why the service is attractive, the techniques which may be used, and selected specific system proposals which have been made. II. The Need for Long Range Communication [5, 6]

At the present time various points around the world are intercon­ nected by a variety of communication techniques. These include radio service and cable service. In long distance communication between con­ tinents, the radio service is in the high frequency band, between approxi­ mately 3 and 30 Mc. The intercontinental cable services are by subma­ rine cable. In the highly developed areas of the world, especially in Europe and the United States of America, long distance continental service is avail­ able. Here the radio service is primarily in the microwave band, around 4000 Mc, over distances of about 30 miles. There is an increasing use of frequencies between 100 and 1000 Mc for scatter propagation transmis­ sion, over distances of several hundred miles. Cable and open wire trans­ mission lines are widely used, but their importance is becoming relatively smaller as the radio techniques develop. The communication relay satellite is of particular importance in the long range service of the intercontinent type. A. Types of Communication Demand

In the intercontinent service, the types of communication traffic which are in use at present or are expected to be used in the near future include: (1) telegraph messages, (2) facsimile, (3) narrow band data transmission, (4) telephone messages, (5) wide band data transmission, and (6) tele-

COMMUNICATION RELAY SATELLITE TECHNIQUES

23

vision relay. All of these are currently in use, but at present wide band data are not commonly encountered. In considering the possibilities of intercontinent relay by satellite, it appears that all of these types of traffic must be accommodated. Fortu­ nately, this is not too difficult, since groups of messages of one type have many of the characteristics of a single message of a more difficult type. For example, a group of telegraph signals occupies the bandwidth of a single telephone signal, and may be readily transmitted over a telephone circuit. In the same way, groups of telephone signals occupy approxi­ mately the same bandwidth as a TV signal and may be transmitted over a TV relay circuit. From this it is observed that adequate flexibility can be provided if the satellite system has sufficient bandwidth, and if this bandwidth can be subdivided to suit the particular type of traffic to be carried. The subdivision can be accomplished at the end terminal, so the major requirement for the satellite is the provision of adequate band­ width. B. Factors Affecting Demand

The demand for rapid long distance communication is determined by many factors which are difficult to evaluate. Some of the important ones are: (1) (2) (3) (4) (5) (6) (7)

Size of population in an area Standard of living Amount of trade and industry Existing communication capabilities Political, military, and habit factors Time difference between areas Differences in language between areas

In considering these factors it is important to distinguish between the potential demand and the current demand. The major difference between these lies in the matters of existing communication capabilities and habit patterns. The difference is great where the development of an area is changing rapidly. For those areas having a long history of industrial de­ velopment, the present capability will be close to the potential demand. However, for an area which is recently developed, the current communi­ cation capacity will be much less than the potential demand. C. Estimation of Future Demand

In attempting to evaluate the future need for communication, fore­ casts may be prepared by projection of existing communication growth

24

R. P. HAVILAND

MESSAGES - THOUSANDS

rates, or they may be prepared by consideration of the basic factors which affect the communication potential. Figure 1, prepared by the American Telephone and Telegraph Company [6], is an example of the first type of prediction. This shows the overseas telephone communication traffic

1920

1930

1940 YEAR

1950

I960

FIG. 1. Overseas telephone messages 1920-1959.

leaving the United States, which is the most important communication customer of the world. From this it is evident that the average growth rate over the past forty years has resulted in a fivefold increase in traffic for each ten-year period. There have been variations in growth rate, and even decreases. The decreases correlate well with major world events such as depressions and wars. Each of the decreases has been followed by a period of relatively rapid growth, some of which correlate with improve­ ments in service. The second method of estimating demand can be developed by assign­ ing relative weights to the various factors, listed above, which govern the

0

/f'

FIG. 2. Demand for communication service : 1970.

SOLID BLACK-UNFILLED DEMAND NUMBER-TOTAL DEMAND NORTH ATLANTIC = 100

SHADED- PRESENT AND SCHEDULED CAPACITY

COMMUNICATION BELAY SATELLITE TECHNIQUES

25

26

R. P. HAVILAND

demand for communication. An example of this type of study, prepared by the General Electric Company [6], is shown in Fig. 2. In this figure the total width of each line represents the potential demand estimate for 1970. The lightly shaded part of the line represents existing circuit ca­ pacity and the black section estimates unfilled demand during 1970. Numerical values shown for each line represent the relative total demand, with the North Atlantic path being taken as 100 units of traffic. It may be noted that the relative demand changes in a rather slow fashion with time, whereas the actual capacity may change rapidly. This is due to the relative slowness with which industrialization and popula­ tion of areas change. These two methods of traffic estimation may be combined to secure an estimate of the total communication demand for the world. For ex­ ample, extrapolation of all types of traffic across the North Atlantic is 6 expected to equal about 6 X 10 message units per year in 1970 or a total 8 for the world of approximately 36 X 10 message units per year. Here a message unit is defined as a signal or a group of signals equivalent to a 3-min telephone conversation. For the period around 1980, total traffic is estimated to amount to between 3 and 5 times the 1970 traffic. III. Satisfaction of Communication Demand [7]

The present and expected future modes of intercontinental (transocean) telecommunications are shown pictorially in Fig. 3. Two of the

FIG. 3 . Modes of telecommunication.

COMMUNICATION RELAY SATELLITE TECHNIQUES

27

modes use radio communication, one being by satellite and the other by ionosphere reflection. The third uses submarine cables. Satisfaction of the demand for long distance communication by means of ionosphere-reflected radio, shown in Fig. 1 and Fig. 2, appears to be virtually impossible. The major reason for this is that the usable high frequency band is so narrow that it is currently overloaded. In addition, the usable band is becoming narrower with time, due to decreasing solar activity, which determines the usable frequency range. Further, the quality of transmission in the H F band is not completely satisfactory, due primarily to disturbances in the ionosphere that cause signal fades and even loss of signal. For these reasons additional communication modes have been sought. The second existing method of communication is the submarine cable. The basic technique is very old, but has recently been revitalized by the development of long life undersea amplifiers. This has greatly improved the quality of transmission and also the communication capacity per cable. Further improvements are expected. Current and expected future repeater cables are technically capable of providing both the communi­ cation quality and capacity required in the future. At the present time high quality cables of the repeater type have been installed across the North Atlantic, and to Alaska, Hawaii, and to some of the Caribbean islands. Operational experience has been quite good, the cables having shown good average reliability and high quality. Plans have been made for an extension of these cable routes, including a plan for a globe-girdling network passing primarily through countries asso­ ciated with the British Commonwealth. The earlier repeater submarine cables had a relatively high installa­ tion cost, on the order of $50,000 per mile. Recent development and im­ provements have lowered this cost, and it is expected that the cost will approach $10,000 per mile in the near future. Even with this reduction the installation cost for a system remains high, approximately $30,000,000 for a North Atlantic cable and on the order of $400,000,000 for the world cable. In addition, the cables do have a limitation in bandwidth. Current cables can handle several hundred kilocycles and it is expected that fu­ ture cables will handle up to 1 Mc. These bandwidths are adequate for telephone and telegraph traffic but are not sufficient for high quality tele­ vision relaying. Such relaying may be achieved by connecting cables in parallel or by transmitting recorded video signals at a reduced rate. Be­ cause of the high cost of such operation, and due to interference to other types of traffic, television relaying by cable has not been common. I t is not expected to be of great importance in the future.

28

ft. t\ tiAViiAttr)

The fact th&t submarine cables can provide high quality service for nearly all the long communication demand indicates that a satellite sys­ tem must meet certain standards if it is to be preferred over the cable system. Basically the satellite system must meet four criteria if it is to be used instead of the cables. These are: (1) The satellite system must provide a bandwidth at least equal and preferably greater than that provided by the submarine cable. Bandwidths of at least 3.5 Mc are needed for television relaying. From 6 to 10 Mc is desired as a minimum to permit simultaneous use for television and other types of traffic. (2) The satellite system must provide a high quality service having good reliability. This means that the signal-to-noise ratio must be high and the outage time very small. (3) The satellite system must interconnect major communication countries of the world and preferably should make possible in­ terconnection of the important population points. This means that the system must provide coverage between approximately 60 deg N latitude and 40 deg S latitude by line-of-sight radio communi­ cation. (4) The satellite system must compete on a cost basis with cables. This cost comparison must be based both on the initial installa­ tion and operation cost of the system and on future capacity for growth. Many studies of the means of meeting these criteria, and of the fac­ tors which affect them have been made and are reported in the literature. The following material will discuss several of the factors involved and then review several of the system proposals which give an integrated world-wide system, capable of meeting the above criteria.

IV. Basic Satellite Communication System Techniques [1-4, 6]

The simpler type of satellite is the passive reflector satellite shown schematically in Fig. 4. Here an Earth transmitter beams energy to a reflector satellite, usually a sphere which reflects the energy in all di­ rections. A part of this energy is picked up by receiving stations. The system has the advantage of great satellite simplicity but extremely high transmitter powers and very large satellites are required if an ap­ preciable bandwidth is to be handled. For example, using 85-ft diameter antennas at the terminals, a 10-kw transmitter can handle two telephone channels (or equivalent) with a 100-ft diameter satellite. Bandwidth

COMMUNICATION RELAY SATELLITE TECHNIQUES

29

could be increased by increasing the size of the antenna, the size of the satellite, or the transmitter power level; but it appears that it would not be possible to provide a system capable of equaling the quality of the submarine cable except at many times the cost of the cable. Therefore,

FIG. 4. The passive reflector satellite.

it seems that the passive reflector system is not satisfactory for this type of service. The second type of system, and the one used in the first communica­ tion satellite experiments, is the delayed repeater shown schematically in Fig. 5. Here, signals are transmitted to the satellite where they are re­ corded. When the satellite moves to the neighborhood of the message destination, the recording is played-out for receipt by the destination station. This system appears to be comparative with the cable in cost and quality but cannot provide the real time service required by tele­ phone traffic, computer traffic, two-way telegraph operation and so on. Therefore, it seems that the system will be used, at most, for special situ­ ations where the long intervals between transmission and receipt are not disadvantageous. The remaining method of system operation is the active satellite re­ peater shown schematically in Fig. 6. Here the signal is received by the

30

R. P. HAVILAND

satellite and immediately retransmitted. The system requires less power than the passive reflector, since the power is proportional to the square of the distance to the ground station, rather than the fourth power as in the passive reflector. This means that a given power level and antenna size

FIG. 5. Delayed communication mode.

can handle a much larger amount of traffic. Since the signal is retrans­ mitted immediately, the system is essentially a real time system, suitable for all modes of traffic. Therefore, it appears that a usable relay system must be based on the active repeater. REPEATER

FIG. 6. Active satellite repeater.

31

COMMUNICATION RELAY SATELLITE TECHNIQUES V. Design Factors in Satellite Systems [8-12] A. Coverage

[3-4]

As is well known, the amount of the surface which is visible from a satellite increases as the satellite altitude increases. Normally, in con­ sidering this coverage, it is only necessary to take into account the amount of Earth surface which is visible from a single satellite. However, in the communication service the problem of coverage is modified, since the establishment of a communication path requires that the satellite be simultaneously visible from both of the Earth terminals. This changes the coverage zone from a circle to an oval shaped figure, which results from the intersection of the two visibility circles of the individual stations. The coverage can be considered in terms of the time of mutual visi­ bility, this being the time during which the satellite is within the mutual coverage zone. Because of the large number of possible combinations, calculation, or even estimation, of the world surface coverage is difficult. The usual procedure involves selecting a group of typical communication terminals and estimating the number of satellites in various orbit planes and altitudes needed to give a required grade of service. General conclu­ sions are then drawn. Over a given route the two most important factors are the latitude of the route midpoint, which gives the location of the mutual coverage zone; and the size of this zone, which is determined by route length and satellite height. For example [3], consider the route from Newfoundland to the Heb­ rides in the British Isles, with the satellite in a 3000-nautical-mile polar orbit. For this route the mutual coverage zone encloses the pole so that a polar orbit satellite will give some service on each pass. For this situa­ tion, using a 3%-deg minimum elevation angle, we have the following: Period of satellite—195.2 min Shortest visibility—21.5 min Longest visibility—26.5 min Average visibility—19.6 per cent For a given average fraction of service interruption (i), the number of satellites (n) is

F = average visibility For the case above, for various values of i i (per cent) n

10 11

1 21

0.1 32

0.01 42

32

R. P. HAVILAND

That is, a rather large number of satellites are needed to give good grade service. Even with the numbers shown there is small but finite risk of a long wait. For example, for the 32-satellite system the chance of a 10-min interruption is 0. 01 per cent. When other paths are considered the minimum number of satellites needed changes [6]. For example, consider a north-south path of the same length as the above, with the midpoint at the equator. For the same height and elevation angle this gives: Period of satellite—195 min Shortest visibility—0 min Longest visibility—41.2 min Average visibility—9.8 per cent The required number of random satellites is i (per cent) 10 1 0.1 0.01 n 22 46 66 88 The number of satellites needed on a given path can be changed by changing the inclination, orbit altitude, and using station keeping, which involves maintaining the satellite in known position with respect to each other. Figure 7 shows the number required for two typical paths with uniUJ

3*1

UJ

32

NEWFOUNDLAND-SCOTLAND

MIAMI-BUENOS AIRES

>

o o 28

O

5 O

EQUATORIAL ORBIT 201

CO IftJ

60° INCLINED ORBIT

CO

UJ CD

2

POLAR ORBIT

EQUATORIAL ORBIT

3

SATELLITE ALTITUDE, NAUTICAL MILES X I O "

3

F i g . 7. Number of satellites needed for continuous service.

COMMUNICATION RELAY SATELLITE TECHNIQUES

33

form satellite spacing. For random spacing the number can be obtained by solving the equation above using the reciprocal of the number of uni­ form-spaced satellites as the average visibility. Table I gives a tabulation for a number of additional paths. TABLE I . NUMBER OF EQUALLY SPACED SATELLITES FOR 100%

H = 3000 miles

Path

Polar

Manila-Honolulu Bombay-Perth New York-Lisbon Lisbon-Moscow Honolulu-Wellington Miami-Buenos Aires Anchorage-Tokyo Newfoundland-Lisbon

64 63 11 9 55 44 14 7

45° 331 43 14 11 36 32 22 10

SERVICE

0

H = 6000 miles

Equa­ tor 25 8 71

b

7 6

b

40

Polar

45°

17 15 6 5 15 15 6 5

13 10 9 7 10 10 10 7

Equa­ tor 10 6 10 10 4 4

b

8

°6 Minimum elevation 7}i°. (—) no direct service.

It is evident from the brief discussion that the number of satellites required varies markedly with altitude, as would be expected. The polar satellites give best coverage away from the equator. They can give cov­ erage to all parts of the world if a sufficient number of satellites is used. In general, the equatorial system shows a smaller variation in the re­ quired number of satellites, but cannot give coverage for those routes having their midpoint near the poles. It is evident from this discussion that there is a preference for the higher altitudes since they give the greatest coverage with a small num­ ber of satellites. However, before concluding that the satellite system should operate at high altitude, it is necessary to consider several addi­ tional factors. The most important of these is the time required for signals to traverse the path, Earth to satellite to Earth.

B. Time Delay [10, 13]

Figure 8 shows the magnitude of the time delay plotted against ground station separation, for several satellite altitudes and for several angular separations between the location of the ground station and the

34

R. P. HAVILAND

plane of the orbit. From this figure, it is evident that the one-way time delay varies by a relatively small amount for a given ground station separation and satellite altitude. For altitudes below about 3000 miles, the coverage from the satellite increases at essentially the same rate as

TIME DELAY, MSEC FIG. 8. Time delay—single hop.

the time delay. At altitudes greater than 6000 miles the time delay con­ tinues to increase but the coverage does not increase appreciably. Where the separation between ground stations is greater than the coverage of the satellite, it becomes necessary to introduce an intermedi­ ate relay station to establish the communication path. This, of course, increases the time delay. This increase is shown in Fig. 9, which shows the minimum time delay (for reasonable location of the additional relay station) plotted against ground station separation. This indicates that time delays for communication paths reaching halfway around the world can become quite long, but in general are largest for the higher altitude systems. To see the importance of this time delay it is necessary to consider characteristics of the communication traffic. Modern telegraph traffic is handled by automatic printers which incorporate error-correction circuits. If an error is made, the receiving station sends a query to the transmitter, which stops transmission, corrects the error and resumes traffic. In this circuit a delay between transmission and receipt of signal has negligible effect. However, if there is a sudden change in the delay, which could be caused by switching traffic from one satellite to another, there will be a gap in the signal or simultaneous reception of two signals. Both will be

COMMUNICATION

RELAY SATELLITE

TECHNIQUES

35

interpreted as an error by the machine. Therefore, for telegraph traffic it is desirable that the change in time delay be zero. This is most easily satisfied by the stationary synchronous orbit, which does not require switching. I t may be satisfied by other systems if the switching is ar810

3 HOPS 19,316 N. Ml.

790 770 750

h

2 HOPS 19,316 N.MI.

530>440

TIME DELAY, MILLISEC

3 HOPS 19,316 N.MI

420 400 2 HOPS 19,316 N.ML

'330

H

310 290

SATELLITE TO GROUNO RELAY**(READ ON LEFT-HAND SCALE)

3 HOPS 6000 N.MI.

- \ 270 250

INTERSATELLITE RELAY^ 'l70 (REAO ON RIGHT-HAND H 150 SCALE) *GROUN0 TERMINALS ARE ON THE EQUATOR *l HOP IS THE SAME FOR BOTH CASES J 1 I I I I i i i

- | 130 110 90

70 12 13 14 15 16 17 16 19 20 21 22 23 24 25 3 N.MI. I0 DISTANCE BETWEEN 6ROUN0 TERMINALS 7 FIG. 9. Time delay—multiple h o p ; relaying.

ranged so that the time delay is identical at the instant of switching. This is relatively easy for the equatorial orbit but becomes more difficult for the polar orbit. In the case of telephone traffic, changes in delay are barely perceived. However, an additional factor must be considered: the presence of echo on the telephone lines. This exists because a single pair of wires are used to carry telephone messages in both directions. In short distance trans­ mission this echo is not bothersome, but in long distance transmission it is perceived, especially by the talker. Under severe conditions, when the echo is strong and the time delay is long, the presence of the delayed echo interferes with the talker's ability to speak, causing him to stammer, to speak more slowly, or even to stop speaking. As a result, it has been found necessary to set international standards for the magnitude of time delay and echo intensity. Unfortunately, these standards do not cover circuits as long as will be used in satellite communication. Therefore,

36

fe. P.

ftAVlLAND

revised standards are needed. Work is progressing on these standards but has not yet reached the point of formal adoption. I t is not possible, at this time, to make a statement as to the limitation which echo and time delay places on satellite height. There is some indication that it will be necessary to limit the time delay to about one-third of a second (350 msec). If this becomes the internationally accepted value, it will mean that the synchronous satellite system will be limited to a single-hop telephone traffic and that the long distance transmission will be carried over several hops by a lower altitude system. This may lead to the installation of two separate systems, or it may mean that the low altitude system will handle all traffic. I t is expected that this matter will be considered at forthcoming in­ ternational conferences and that eventually standards will be adopted and the preferred communication system established. C. Frequency Allocations and Sharing [6, 14-16]

At the present time, by international agreement, the radio frequency spectrum is divided into a group of bands. These bands are allocated to one or more types of communication service. The existing services have made use of the entire spectrum. In many cases the occupancy or density of use is becoming quite high. In general, this occupancy is greatest at the lower frequencies, since these were the first developed and since equipment is generally less expensive if the frequency can be kept low. At present none of the allocations authorize the use of satellites for the type of traffic to be carried by relay satellites. Consequently, some methods for accommodating this new service must be found. One method of providing allocations is to displace some existing com­ munication service and substitute an allocation for satellite service. This is simple in principle, but is very difficult in a practical sense. Space must be found for the displaced services and undue burden may be placed on such a service by the necessity of purchasing and installing new trans­ mitters, receivers, and antennas for the new frequency. Such displacement has occurred in the past, for example, to provide frequencies for space research, but the international agencies are understandably reluctant to adopt this procedure. The second possible method of securing allocations is to impress the new service on top of an existing service, that is, to establish a shared allocation. The criteria for such sharing obviously must include the fac­ tors that the new service not cause harmful interference to the original service and that the original service not degrade the performance of the new service unreasonably.

COMMUNICATION RELAY SATELLITE TECHNIQUES

37

In addition to these factors, attention must also be given to natural phenomena that affect the new service. In the case of satellite transmis­ sion, the phenomena of primary importance are reflection and absorption by the ionosphere, and absorption by the atmosphere. These establish a lower limit of possible allocations of about 100 Mc and an upper limit of about 10,000 Mc. When the available frequency bands within this range are examined, it is found that the bands below about 2000 Mc are so heavily occupied that sharing does not seem feasible. A number of the remaining bands are not available; some are utilized for safety services, where any interfer­ ence is dangerous, and others allow pulse emission (radar location), and have undesirable interference properties. Of the remaining bands, the most attractive for sharing are those which are assigned to the point-topoint service and used for microwave relay. Such relays are now used over the entire world to carry telephone, telegraph and TV signals of the same type as would be carried by the satellite system. This relay service uses relatively low power transmitters with highly directive antennas aimed at another relay station and accordingly do not cause much inter­ ference. Nor are they particularly susceptible to interference. The one problem which appears is a potential interference to the relay system by the ground transmitter used in the satellite system. However, it is likely that interference can be reduced to an acceptable level, or eliminated, by proper choice of the site of the satellite ground transmitter. In many such locations advantage can also be gained from the natural terrain by using intervening hills to provide shielding. At this time there appears to be general agreement that this type of sharing should be adopted on an international basis. The United States, in particular, has issued a preliminary proposal which would allocate about 3000 Mc of bandwidth on this shared basis. The matter will be considered by an international conference late in 1963 and it is likely that the shar­ ing proposals will be adopted. D. Attitude Control and Station Keeping [6, 7]

Attitude control of the communication satellite is desirable for two reasons. First, it permits the antenna to concentrate the signal on the usable area of the Earth, thereby reducing the required transmitter power for a given signal quality. For example, in the synchronous orbit, attitude control permits use of an antenna having a beam width of 17 deg, which reduces the required power by a factor of about 100. The effect is less important at lower altitudes, but even in the 3000-mile orbit power reduc­ tions of about 20 are possible. The second reason for stabilizing the satel-

38

R. P. HAVILAND

lite (and the antenna) arise from considerations of constancy of signal. Practical antennas never radiate equally in all directions, and if the an­ tenna is moving a given signal will show variation in signal strength which must be corrected. In addition, rotation of the antenna will intro­ duce a small frequency shift, which may be important in some systems. The term "station keeping" is used to designate satellites which main­ tain a specified position, either with respect to other satellites, or with respect to a point on the Earth's surface, as is necessary for stationary satellites. Such station keeping demands reasonably close control of the altitude of the satellite and the inclination of its orbit, and extremely close control of its velocity. This control is obtained by adjusting the velocity of the satellite by a small rocket motor. This, in turn, requires an attitude control system to insure that the thrust of the motor is in the required direction. Therefore, station-keeping satellites always require an attitude control system. The two major lines of current attitude control effort involve the use of so-called passive stabilization and the use of horizon sensors. Some work is being done on other systems; for example, on systems using gyro elements which interact with the motion of the satellite about the Earth to give the required attitude control. The passive stabilization system depends on the fact that the gravita­ tional attraction of the Earth decreases with altitude. If the satellite is symmetrical about an axis of rotation, then the part which is closest to the Earth will be attracted most strongly. The satellite consequently will try to align itself to the vertical, thus maintaining the same face towards the Earth just as does the Moon. While attractive in principle, the sysFILL AND VENT

ATTITUDE SENSOR

„ PNEUMATIC $ MANIFOLD VALVE

STORED COLO H2

REGULATOR

DATA PROCESSING

ATTITUDE CONTROL AMPLIFIER

c NOZZLE

MOMENTUM PACKAGE

FIG. 10. Attitude control subsystem block diagram.

39

COMMUNICATION RELAY SATELLITE TECHNIQUES

tern has not yet been reduced to practice. The chief reasons for this lies in the extreme smallness of the force involved (which requires a very great accuracy in symmetry) and in the need for providing damping forces to stop rotation and swinging of the satellite. The horizon stabilization system depends on a measurement of tem­ perature at the interface between the warm Earth and cold space. If the satellite is in its correct position the signal from two opposite parts of the horizon will be equal. However, if the satellite makes an angle with the local horizontal, one temperature sensor will see a larger fraction of the Earth than the other. This may be used to generate a signal to cor­ rect the attitude of the satellite, either by a small rocket jet or by chang­ ing the velocity of a flywheel, which changes the altitude through momen­ tum conservation. An elementary block diagram of a single axis attitude control system [6] is shown in Fig. 10. This uses stored nitrogen gas to provide reaction forces and also uses a flywheel for momentum control. The flywheel serves for small, slow speed corrections and the jets for larger and more rapid corrections. The system is intended for use with one of several forms of horizon sensors. A minimum control system would use two channels, one in pitch and a second in yaw. A third axis of control would be required if roll stabilization is needed. Such stabilization might be with respect to the Sun. The orbit control system [6] shown schematically in Fig. 11 also uses FILL AND VENT

PRESSURIZED TANK BLAOOER

ST OREO COLO N,

FILL AND VENT

FUEL MANIFOLD

COMMAND RECEIVER

NOZZLE AND THRUST PROGRAM

I I

0J1 H&h™ ANOZZLES

FIG. 11. Orbit control subsystems block diagram.

40

R. P. HAVILAND

stored nitrogen gas to change the velocity of the satellite, thereby correct­ ing its orbit. This system is designed for control from the ground, the necessary orbit measurements and calculations being accomplished by ground equipment. Obviously, proper functioning for this system depends on the presence of an attitude control system. If directive antennas are used on the satellite the attitude control system must remain in operation for the life of the satellite. This poses a considerable problem in reliable design. I t appears that the orbit control system does not need to have as great a life since calculations indicate that the orbit can be brought to the desired accuracy in a matter of a few weeks. Thereafter, there would be some variation due to atmospheric drag and perturbations by the Earth and other bodies, but it appears that these changes are minor and can be easily compensated by the ground equipment. E. Power Supply [6, 7]

Up to the present, all power supplies proposed for communication satellites are based on the absorption of solar energy by silicon solar cells. Batteries are also provided to supply energy during the time the satellite is in the shadow of the Earth. The solar cells must also charge these bat­ teries while the vehicle is in sunlight. An elementary diagram of such a solar power system [6] is shown in Fig. 12.

SOLAR ARRAY

SOLAR RADIATION

SOLAR ARRAY AND BATTERY CHARGE CONTROL

VEHICLE EQUIPMENT POWER UNREGULATED

VEHICLE EQUIPMENT POWER

B

REGULATED EQUIPMENT POWER SUPPLY

VEHICLE EQUIPMENT POWER

REGULATED

FIG. 12. Block diagram of typical solar power system.

BATTERY;

COMMUNICATION RELAY SATELLITE TECHNIQUES

41

Two different approaches for mounting the solar cells are being used in current and proposed designs. In one, the solar cells are mounted over the entire vehicle body. For this system the solar charge control consists primarily of small diode elements which effectively disconnect the cells shadowed by the vehicle body. In the second system the solar cells are mounted on paddles which are oriented to face the Sun. This is accom­ plished by using the roll axis of the vehicle to turn the paddles toward the Sun and by rotating the paddles so that they are always normal to sunlight. This system has the advantage of greater absorption efficiency for solar energy and requires fewer solar cells but does require active orientation elements in the satellite. One of the major problems in the design of both types of systems is the battery. Considerable problems have been encountered in securing satisfactory seals against the vacuum of space and in preventing internal shorts in the cells of the batteries. An additional source of power which may be used in future designs is the radioisotope power supply. A typical design uses the energy re­ leased by a radioactive isotope such as plutonium-238. This thermal energy is converted into electrical energy by thermocouples. Such a supply has the advantage that energy is continuously available, which eliminates the need for storage batteries. However, there is a problem in handling such equipment due to the radiation from the radioactive source. In addition, the cost of these units, at the power level required, has been relatively high. These factors have prevented use of the system in com­ munication satellites to date. VI. Specific System Proposals [6,

17]

A number of proposals have been made for a specific satellite relay system. Each of these has combined the various possible parameters of the system in different ways, the goal being an integrated system meeting the designers' objectives. In the following paragraphs three of these proposals have been se­ lected to show typical major features of the systems. For convenience, the basic system characteristics are presented first in short descriptive form, and then are given in tabular form. Finally, the relative advantages and disadvantages are listed. The three system proposals chosen [17] are typical of systems based on low, intermediate and synchronous orbits. The low altitude system would use 50 satellites in random polar or near polar orbits at an altitude of 2500 miles. The satellites would be designed for simplicity, using omnidirectional antennas and solar absorb­ ers and would be of low enough weight to permit several satellites to be

42

R. P. HAVILAND

carried by one carrier vehicle. Long life, about ten years, is expected. High quality service is provided to any point on Earth. The intermediate altitude system would use ten satellites in an equatorial ring at an altitude of 6000 miles. Attitude control and station keeping are used, with an oriented solar absorber. Because of the added complexity of these elements, weight would be greater than the first system, and life would be from three to five years. The system is intended to handle a large amount of high quality traffic. The synchronous (stationary) satellite system is intended to be launched with a small carrier vehicle, and so is designed for minimum weight. It is also intended that the system be installed in stages, making use of the ability of a single stationary satellite to cover about one-third of the Earth's surface. The details of the three systems are shown in Table II. The comparative advantages and disadvantages of these systems have been stated [17] to be: (1) The low altitude system covers the globe. The equatorial systems give coverage only to about 60 deg and 70 deg latitude for the intermediate and stationary systems. (2) The stationary system can give continuous coverage to selected stations with a single satellite and full coverage with three oper­ ating satellites. The low altitude systems require a larger number, up to 50 for the lowest altitude system. (3) The low altitude system does not need the precise control of orbit required by the equatorial systems. (4) The synchronous system does not require moving antennas, nor does it have doppler shift and traffic switch-over problems. These factors become more marked as the altitude is reduced. (5) The synchronous system uses fixed antennas which reduce inter­ ference problems. The intermediate altitude system antennas move, but reach a low elevation angle at only two azimuths. The low altitude system antennas must look in all directions, and so have the most intensified interference problem. (6) In the synchronous system many ground stations operate with a single satellite. As a result the total frequency band for a given amount of traffic is larger than that for other systems, which "re-use" the same frequencies in different parts of the world. (7) The time delay in the synchronous system exceeds current speci­ fications. As can be seen from these comparisons, the importance placed on a given factor or combination of factors will lead to a preference for one

COMMUNICATION RELAY SATELLITE TECHNIQUES

43

TABLE I I . PROPOSED SATELLITE RELAY SYSTEMS

Low system General Altitude Orbit Number of satellites

Intermediate system

Synchronous system

2500 miles Polar random 50

6000 miles Equatorial spaced 10

22,300 miles Equatorial spaced 1 + spare (6 world wide)

1 kw two-60-ft dishes two-50-ft horns

Less than 10 kw two-60-ft dishes Same

2800 watts 150-ft dishes Same

Unknown (low) 7 deg

100°K 7 / 2 deg

80°K 10 deg typical

28 typical

18 initial: up to 130

2

1 watt Isotropic Same

Less than 10 watts 53 deg Same

2.5 watts 18 de? Dipole

Weight Estimated life

150 1b 10 years

2900°K Paddle mount solar cells 500 watts Horizon scan Rectangular and paddles 1000 lb maximum 3-5 years

1250°K Solar cells

Primary power Stabilization Shape and size

3000°K Surface mount solar cells Unknown Spin 4-ft sphere

15 watts Unknown Cylinder 28 in. 0 X 24 in. 33 1b 1 year

Unknown

3 from one Atlas

Scout

Frequency

6000 Mc

2000 Mc

Bandwidth

125 Mc

6.5 Mc

Modulation Capacity (4 kc Ch.) One-way delay

Large index FM 600 duplex 0.05 sec

Large index FM 600 duplex 0.15 sec maximum

2000 Mc down: 470 Mc down 2.5 Mc up: 25 Mc down SSB up: FM down 300 duplex 0.3 sec

Ground Stations Transmit power Transmit antenna Receive antenna Receive noise temperature Minimum elevation Number stations served Satellites Transmit power Transmit antenna Receive antenna Receive noise temperature Power source

Carrier Vehicle Communications

44

R. P. HAVILAND

or the other of the possible systems. As noted previously, many of these are still being studied. As a result, there is as yet no agreed-upon preferred system. Agreements are expected, and will be reported in the technical journals, which should be consulted for the latest developments. VII. Status of Experimental Communication Satellite Programs

[18]

Two experimental satellite programs have been completed, and sev­ eral others are in progress. A. Completed Programs

The completed programs are Score [19] and Echo 1 [20]. Score was an active repeater, primarily designed to operate in the delayed mode, with received signals recorded on magnetic tape and transmitted on command, when near the desired Earth terminal. Echo 1 was a passive satellite formed by an aluminum-coated mylar sphere 33 meters in di­ ameter. This satellite was still in orbit in 1963, although it had lost its spherical shape. An additional program, now inactive, is West Ford. This passive sys­ tem was intended to use a large number of tuned dipoles as reflectors. The first test of this system failed, since the dipoles did not separate, B. Programs in Progress

The current passive satellite programs are Echo 2 and Rebound [19]. Echo 2 will be a sphere 41 meters in diameter. It is constructed of an aluminum-mylar-aluminum laminate, designed to hold spherical shape after the pressurizing gas lost. This program has completed sub­ orbital test, and should be in orbit during 1963. Rebound will carry three of these spheres, placing them in a circular orbit at preselected intervals. Advent is a military communication satellite system. Originally de­ signed for the synchronous orbit, the program was fully redirected in 1963. I t is expected that the redefined program will include a low alti­ tude system and a synchronous system. Project Relay, now in development, uses a spin stabilized platform in an inclined elliptical orbit. Two repeaters are carried, each capable of relaying a television channel, 300 one-way telephone channels or 12 twoway telephone channels. The first Relay was launched on December 13, 1962. Project Telstar [21], now operating, also uses a spin stabilized plat-

communication

relay

satellite

techniques

45

form in an inclined elliptical orbit. The satellite carries one active re­ peater capable of transmitting a 3-Mc band. This has been used to relay television, voice, and other signals between the U.S.A. and Europe. Project Syncom, also in development, is intended for the 24-hr inclined orbit. The active repeater is designed to relay a single two-way telephone signal. The satellite is spin stabilized, and includes a station keeping system. Syncom 1 was launched on February 14,1963. Ground stations for these satellite programs have been or are being installed in a number of locations around the world. These include horn antennas at Andover, Maine; Homdel, New Jersey; and in France. There are parabolic dish antennas at Goldstone, California; Monmouth, New Jersey; and in England, Brazil, and Japan, as well as other locations. In addition, antenna and receiving systems designed for other purposes, such as radio astronomy, have been used for various communication experi­ ments. The experimental results have been very good. System performance has been in agreement with calculations, and propagation has been very close to theoretical values [22], generally within the limits of experi­ mental error. A wide variety of traffic has been transmitted on a scheduled basis, including relayed programs which have been watched on millions of home receivers [23]. As a result, the relay satellite principle can be considered to be fully demonstrated. References 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13.

Clarke, A. C , Extra-terrestrial relays. Wireless World 51, 10, 305-308 (1945). Haviland, R. P., Can we build a station in space? Flying 51, 10, 45 (1949). Pierce, J. R., Orbital radio relays. Jet Propulsion 25, 153-157 (1955). Haviland, R. P., The communication satellite. Astronaut. Acta 4 (1958). Federal Communications Commission, "Statistics of Communication Common Carriers" (Annual). Government Printing Office, Washington, D. C. Material Filed with the Federal Communications Commission in Reply to FCC Dockets 13522 and 14024: FCC Reading Room, Washington, D . C. "Communication Satellites," Hearings before the Committee on Science and Astronautics, Government Printing Office, Washington, D. C. Brown, R. H., and Lovell, A. B. C , "The Exploration of Space by Radio." Chapman & Hall, London, 1957. L. J. Carter, ed., "Communication Satellites." Academic Press, New York, 1962. Convention Record, 5th Natl. Symposium on Global Communications, Chicago (available from C. F. Wittkop, 1450 N. Cicero Avenue, Chicago, Illinois). Space Electronics Issue, Proc. I.R.E. 48, 4 (1960). Pierce, J. R., and Cutler, C. C , Interplanetary communications. Advances in Space Sci. 1, 55-109 (1959). Secretariat of the C.C.I.T.T., Analyses and conclusions from documentation within the C.C.I.T.T. about the influence of propagation times and of echo ef-

46

14.

15.

16. 17. 18.

19. 20. 21. 22. 23.

r.

p.

haviland

fects on the telephone transmission quality of international calls, Document IV-55, International Radio Consultative Committee, Geneva, March 9, 1962. "Preliminary Views of the United States of America, Frequency Allocations for Space Radiocommunication." Federal Communications Commission, September 7, 1961. Draft Report, Frequency sharing between communication satellite systems and terrestrial radio services, Document IV-79, Study Group IV, International Radio Consultative Committee, Geneva, March 19, 1962. Curtis, H. E., Interference between satellite communication systems and com­ mon carrier surface systems. Bell System Tech. J. 61, 3 (1962). "Frequency Allocations for Space Communications," Joint Technical Advisory Committee, Inst. Radio Engineers and Electronic Inds. Assoc., March 1961. United States of America, Report on the experimental communication satellite program in the United States of America, Document IV-56, International Radio Consultative Committee, Geneva, March 12, 1962. Brown, S. P., and Senn, G. F., Project Score. Proc. LRU. 48, 624-630 (1960). Project Echo, Bell System Tech. J. 60, 4 (1961) (Special issue devoted to re­ sults from Echo). Shenum, R. H., Telstar satellite, and results of the first Telstar experiments. Presented at 13th Intern. Astronaut. Congr., Varna, Bulgaria, September 1962. Haydon, G. W., Optimum frequencies for outer space communications. J. Re­ search Natl. Bur. Standards (Radio Propagation) 64D, 105-109 (1960). Bray, W. J., The world-wide relaying of television by artificial Earth satellites. Presented at Intern. Television Convention (I.E.E.) London, November 1962. Bibliography

"Frequency Allocations for Space Communications," Appendix II. Joint Tech. Ad­ visory Committee, Inst. Radio Engineers, and Electronics Inds. Assoc., March 1961.

Solid Propellant Rocket Technology H.

Thiokol

Chemical

W.

RlTCHEY

Corporation, Rocket Ogden, Utah

Operations

Center

AND

J . M . MCDERMOTT

Thiokol

Development Laboratories, Chemical Corporation, Wasatch Brigham City, Utah

I. Introduction A. Types of Solid Fuel Rockets II. Factors Affecting Propellant Performance A. Composition B. Burning Rate C. Area Ratio D . Mass Rates E. Temperature Coefficient III. Grain Design A. End Burners B. Circular Perforations C. Star Perforations D. Thrust Programs E. Port-to-Throat Ratio F. Mass Ratio IV. Ignition of Solid Propellant Rockets A. Pyrotechnic Igniters B. Pyrogens C. Ignition in Space V. Testing of Solid Propellant Rockets A. Static Tests B. Environmental Tests C. Storage Tests D . Drop Tests E. Detonation Tests VI. Solid Propellant Processing A. Extruded Double Base Propellant B. Cast Double Base Propellant C. Composite Rubber Base Propellant 47

Division

48 48 52 524 55 55 56 57 57 57 58 60 61 62 63 63 64 65 65 65 66 66 66 67 68 68 69 70

48

H. W. RITCHEY AND J . M. MCDERMOTT

VII. Inspection and Quality Control A. Raw Materials B. In-Process Controls C. Cured Propellant D. Radiographic Inspection VIII. Grain Defects IX. Thrust Vector Control and Thrust Termination X. Solid Fuel Rockets for Space Missions XI. Conclusion Bibliography

72 72 73 73 74 74 75 79 85 85

I. Introduction

The principles of rocketry are the same for solid fuel rockets as for large liquid-fuel vehicles. The same laws of thermodynamics apply in both types of motor to the generation of a mass of high temperature gas. Fluid dynamics governs the acceleration of this mass to high velocity. Aerodynamic principles are the same. However, a large and growing technology has developed around solid fuels which takes advantage of the ways in which solids differ from liquids. The solid fuels are more dense and are nonvolatile. They can be stored for indefinite periods in a state of instant readiness. Solid fuel motors can be made extremely large, capable of enormous thrust but with a small number of simple individual components. They may have very high thrust-to-weight ratios or may be so designed that accelerations are well within the endurance of men and instruments. They are receiving more and more consideration for use as motors for space carrier vehicles. This chapter deals with the present state of the art in solid fuel rockets and the ways in which they are de­ signed and manufactured. A. Types of Solid Fuel Rockets 1. B Y MISSION

Solid fuel rockets vary in size from a few pounds to many tons (Fig. 1). They have practically preempted the field of short range rocket and missile artillery involving air-to-air, air-to-ground, and ground-to-air applications. Being relatively cheap to construct, a large number of small rockets can be salvoed or ripple-fired to saturate an area of ground or sky. Free flight missiles without guidance are used for such missions. Rockets of this type frequently have thrust to weight ratios of 100:1. Rockets with greater range use a guidance system to direct them to the target. Well known examples of guided missiles powered by solid fuel motors are the Falcon, Sidewinder, Nike Hercules, and Nike Zeus. The

FIG. 1. Typical solid propellant motor.

UJ

~

o t'4 oo

~

~

a

t%J

~

~

t%J

~

a

~

~

~

>

~

t%J t'1

~

~

o

~

~

t'1 ......

o

50

H. W. RITCHEY AND J . M. MCDERMOTT

relative cheapness of these motors is very important for missiles produced by the thousands or even by the hundreds of thousands. The reliability, storability, instant readiness, and high-launch velocity of the solid fuel motors are the factors which fit these missiles for air-to-air and anti­ aircraft missions. Reliability is of the utmost importance when fired from manned aircraft where a malfunction could destroy plane and pilot. In­ stant readiness and high launch velocity come into play most strikingly at the supersonic speeds of today's aircraft when a split second at launch time makes the difference between a kill or a miss of miles. 2. B Y FUEL TYPE

The solid fuels for rocket motors fall into two broad classes, double base and composites. The double base fuels are compounded of nitro­ cellulose and glycerol trinitrate (nitroglycerine) with stabilizers and ballistic modifiers. A well known example of this type of fuel is JPN. The glycerol trinitrate plasticizes and swells the nitrocellulose and the mass sets to a hard plastic of relatively high strength and low elongation. By methods to be described later, the double base propellant is formed in a mold into a charge (called a grain) sized to fit the rocket case. It is removed from the mold after the propellant sets or cures and the outside is covered with a fire-proof plastic to restrict burning to the central cavity. The inhibited grain is then slipped into the motor case. I t is restrained from longitudinal motion during transportation and handling by springs, plates and washers. The spring-held plate at the head end seals against free passage of gas and is known as an obturator. The aft plate, perforated to make passages for the gas, is called a trap. This method of motor preparation is called cartridge loading. Cartridge-loaded motors are penalized in performance by the weight of the obturators and traps. The case also must be of thick wall construc­ tion, since it is exposed to the hot gas over the entire inside surface and loses strength as it gets hot. The design, therefore, cannot be based on the cold strength of the steel. These limitations restrict the use of car­ tridge loaded motors both as to size and mission. Recent advances in the technology of double base propellants have made possible their use in case bonded designs, eliminating the restrictions on cartridge loaded motors. The composite propellants are made of a fuel and an oxidizer. The oxidizer is an inorganic perchlorate or nitrate. The one most used because of its high performance is ammonium perchlorate. Many materials have been used as the fuel which also serve as a binder for the oxidizer. The important requirement for a fuel is a high heating value. Considered as a

SOLID PROPELLANT ROCKET

TECHNOLOGY

51

binder, mechanical properties are of prime importance. The fuels most used for composite propellants are not themselves explosives. The first fuel binder for composite propellants was asphalt. When plasticized with oil, it could be mixed with the oxidizer but it softened and sagged in hot weather and shrank and cracked in the cold. Among the first polymers to be used were the polysulfide rubbers. Other rubbery polymers have been used both experimentally and in large scale produc­ tion and several more recent ones have shown performance superior to the polysulfides. All that have found extensive use have been liquid polymers that convert with a catalyst or curing agent and heat to a rubbery material. Shortly after World War II the idea was conceived at Jet Propulsion Laboratory under an Army contract that a propellant might be developed that could be poured directly into the motor case, around a mandrel to give the central perforation the proper shape. The propellant when cured in place would be bonded to the case wall. The great advantages of this method of manufacture would be the elimination of the extra weight of the obturators and traps and protection of the case from the hot com­ bustion products. Propellants burn normal to the surface. Therefore, since burning progresses from the central cavity to the wall, most of the case is covered with propellant until burnout and is not touched by the hot gas. Since heat transfer through the propellant is slow, most of the wall remains cool and motors can be designed to the cold strength of the steel, allowing great savings in weight. Very high mechanical strength in the propellant is not required because the steel case provides support. This idea was reduced to practice with rubbery polymeric binders. This was a big advance in fuel technology but it imposed severe requirements on the binder. The coefficient of thermal expansion for most 5 propellants is about 5 X 10~ in./in./°F. For most metals, it runs 3-7 X 6 10~ . Therefore, the propellant must be capable of elongating enough to endure, without cracking, severe stresses from changes in temperature. It must also bond very tenaciously to the case wall; otherwise separation will occur at low temperature. Cracking or bond failure would expose unwanted surface for burning. The propellant must have sufficient resist­ ance to creep so that it will not deform under its own weight. Its mechan­ ical properties must not change drastically on aging. These requirements have all been met, making possible the very large solid fuel rocket motors of today. The use of solid fuel rocket motors to power space carrier vehicles presents few problems that have not already been attacked and solved in smaller rockets for military use. The size of the vehicle is the chief

52

H . W. R I T C H E Y AND

J.

M.

MCDERMOTT

difference. Considerable background in design, materials, and manufactur­ ing methods is required for giant motors. Therefore, the subject of solid fuel rockets in space applications will be deferred until after a discussion of technology that goes into the motors and propellants. II. Factors Affecting Propellant Performance A. Composition

The factors considered important in the ballistic performance of solid propellants are burning rate, specific impulse, volume impulse, pressure exponent, and temperature coefficient of burning rate. All depend on the composition. Composition is to be understood as meaning both the type of material and the weight per cent of ingredients (Table I ) . Both nitrocellulose and glycerol trinitrate, the chief ingredients of double base propellants, are themselves very energetic materials. The energy required for their decomposition is low. They contain within the molecule oxygen for the combustion of the carbon and hydrogen, resulting in a high heating value. In double base propellants, both the burning rate and the impulse depend on the amount of glycerol trinitrate used. The need for improved mechanical properties or lower burning rates has frequently dictated the replacement of part of the glycerol trinitrate with nonexplosive plasticizers such as dimethyl phthalate. The burning rates of double base propellants vary from about 0.2 to 0.8 in./sec at 1000 psi. Specific impulse varies, though not in the same ratio, as more or less explosive plasticizer is used. The fuel binder used in a composite propellant has much to do with burning rate and specific impulse. Some binders burn slowly and some rapidly, depending on chemical structure. The propellant specific impulse is also affected by the structure of the binder. Large amounts of carbon and hydrogen are desirable with little or no oxygen, chlorine, nitrogen, or sulfur. Unfortunately, most liquid polymers which can be cured by convenient methods contain some of these less desirable elements. The amount of oxidizer usually is more important than the chemical structure of the binder on such qualities as burning rate, specific impulse and density. With any chosen binder, both impulse and burning rate go up as the amount of oxidizer is increased. There are practical limits above and below which it is not advisable to go. Propellants with too great a deficiency in available oxygen burn unstably (chuffing). The ideal upper limit of oxidizer would be the amount needed to burn all the carbon to the most efficient mixture of carbon dioxide and carbon monox­ ide and all the hydrogen to water. The practical limit is the point at

Burning rate exponent, n Burning rate at 1000 psia and 70°F, in./sec Specific impulse @ 1000 psia and sea level, lb-sec/lb Density 3 (average), lb/in. Temperature coefficient of pressure, %°F

185-240

0.055-0.62 0.1-0.25

160-220

0.057 0.1-0.8

205-230

0.058 0.1-O.8

Polymer + Curing Agent 16-301 Ammonium Perchlorate 56-78 \ per Miscellaneous 2-14 J cent 0.2-0.5 0.15-1.0

45-55 ] 25-40 ! per 12-22 [cent 1-2 J

Typical composite

0.1-0.8 0.2-0.4

Nitrocellulose Glycerol Trinitrate Plasticizer Miscellaneous

Cast double base

0.1-0.8 0.6-0.9

Nitrocellulose 50-60] Glycerol Trinitrate 30-45 per Miscellaneous 1-10 cent

Extruded double base

COMPOSITION AND PROPERTIES OF TYPICAL SOLID PROPELLANTS

TABLE I . SOLID PROPELLANT ROCKET TECHNOLOGY 53

54

H . W. R I T C H E Y AND J .

M.

MCDERMOTT

which the propellant loses its flow properties and cannot be cast into motors. The density of most of the usable liquid polymers runs between 0.9 3 3 and 1.3 gm/cm . The density of ammonium perchlorate is 1.95 gm/cm . The lower the density of the polymer is, the larger its volume per unit of weight and the more oxidizer that can be used in a propellant. Not as much oxidizer can be incorporated into the high density polymers. This leads to the anomaly that a propellant with high specific impulse (lb3 sec/lb) may have a lower volume impulse (lb-sec/in. ) than a propellant with a lower specific impulse but a higher density. Since burning rate, impulse and density all depend on the amount of oxidizer, it usually is not possible to get very low burning rates (0.2 in./sec) without a sacrifice of impulse and density. Burning rates in the range 0.7-1.0 in./sec are usually achieved with burning rate catalysts. The catalysts most commonly employed are compounds of iron, copper, or chromium. Metal powders, used extensively in solid propellants, provide very high reaction temperatures. Less oxidizer is required because the water in the exhaust is reduced to hydrogen, which is a more efficient working fluid because of its low molecular weight. Metal powders result in an increase in propellant specific impulse and density. There is some accompanying loss of nozzle efficiency, higher with some metals than others. B. Burning Rate

The burning of solid propellants always progresses normal to the surface. The burning rate, which is the rate of regression of the surface, is then measured in inches per second. When sufficient energy is supplied to a propellant surface, decomposition into gaseous products occurs. The gases mix and react, giving off heat. Some of the heat is fed back to the surface to sustain the decomposition. Solid propellants burn faster at higher pressures. One explanation for this is that higher pressure com­ presses the boundary layer between the gas and solid and increases the heat flow to the solid surface. The increased heat increases the rate of decomposition and hence the rate of regression of the surface. In spite of intensive study no completely encompassing theory of burning has been proposed. The burning rate expressions used are inexact and many propellants deviate to a greater or lesser degree. The equation most used is n

r = aPc

(1)

where a and n are characteristic constants and Pc is chamber pressure.

SOLID PROPELLANT ROCKET TECHNOLOGY

55

The exponent n varies from 0.2 to 0.8. I t is most desirable to have a low value. Propellants with exponents about 0.8 are very pressure sensitive which leads to greater difficulty in stabilizing chamber pressure and increased danger of blowing up the rocket from slight imperfections in the grain. For ease of plotting and reading, the burning rate expression is usually converted into linear form log r = J b g c f + n log PC

(2)

A number of other equations have been proposed but no general equation fitting all cases has been found. C. Area Ratio

From what has been said about burning of solid propellants as a surface reaction it is obvious that the larger the surface the greater the amount of gas generated, all other variables omitted. The area of the nozzle controls the rate at which the gas escapes from the chamber. Therefore, chamber pressure is related to these two areas by the equation

P c = b(A /A r B

T

(3)

where b and N are characteristic of the propellant, AB is the burning area, and AT is the throat area. This expression is most used in its linear form. The value of N is greater than 1.0 and is related to n in the burning rate equation as (4) Consideration of this relationship emphasizes why it is desirable to have a low value of n in the burning rate equation. The extreme pressure sensi­ tivity at large values of N imposes the necessity for extra care in the control of composition, throat dimensions and freedom from grain defects. The ratio AB/AT is referred to as KN. D. Mass Rates

The mass burning rate of a solid propellant is (5) where p is propellant density. The mass discharge rate through the nozzle is rhB = ABr

p

mD = ATPCCD

(6)

CD, the discharge coefficient, is a proportionality constant which takes

56

H . W. R I T C H E Y AND J . M . MCDERMOTT

into consideration deviation from the ideal gas laws. As used here, it converts the values to units of lb/sec. The mass burning rate and the mass discharge rate are plotted against pressure in Fig. 2. From this, it is apparent that for any given propellant there is only one stable oper­

FIG. 2. Mass rates vs chamber pressure.

ation pressure for any ratio of burning area to throat area. To the left of the intersection of the two lines gas is being generated faster than it can escape. Pressure rises until the case bursts or the pressure stabilizes. To the right, gas escapes faster than it is being generated and chuffing or intermittent burning may result, or the pressure falls rapidly to the stable operating valve. The rocket designer selects the operating pressure de­ sired and fixes the burning area and the throat area from an experi­ mentally determined curve of PC versus Kn.

E. Temperature Coefficient

The temperature at which a solid fuel rocket is fired affects the cham­ ber pressure. Usually designed for a certain pressure at normal tempera­ ture, the motor develops higher pressure, hence higher thrust, at elevated temperature and lower pressure at lower temperature. The temperature coefficient of pressure is usually expressed as the percentage change of the chamber pressure at normal temperature per degree Fahrenheit change in firing temperature. The values for solid fuels vary from 0.1 to 0.8 per

57

SOLID PROPELLANT ROCKET TECHNOLOGY

cent. The total impulse is only slightly affected, since at lower or higher pressures burning time is shorter or longer and the thrust-time integral remains nearly the same. However, launch acceleration and range are affected. The most useful temperature coefficient is known as 7rk, and is desig­ nated mathematically by the expression (7)

thus providing a measure of the change of pressure with a change of propellant temperature. A second coefficient wp is designed to give the change in burning rate with temperature as measured in a constant pressure environment (strand burning rate chamber) (8)

III. Grain Design A. End Burners

A simple solid fuel rocket might be depicted as a tube completely filled with propellant and burning only on the surface at the nozzle end. The case would then contain the greatest possible weight of propellant. However, the burning area is limited by the diameter of the case and because of the Kn ratio a small nozzle throat area would be required. The thrust of the rocket is described by the relationship F = PCATCF

(where Pc = chamber pressure, A T = nozzle throat area, CF coefficient) or by the equation F = 32.27 s pw

(9) =

the thrust (10)

where 7 s p = propellant specific impulse, m — mass rate of exhaust which is the same as propellant burning rate times density times burning sur­ face. Burning surface area, density, and burning rates usually bear such a relationship that the thrust/weight ratio of end burners cannot match many of the important requirements. B. Circular Perforations

The burning area is increased by making a longitudinal channel or perforation through the center of the cylinder. Burning on the outer sur-

58

H . W. RITCHEY AND

J.

M.

MCDERMOTT

face of the cylinder is prevented by a coating of inhibitor or by case bonding. The grain may be allowed to burn on one or both ends. The perforated grain reduces the amount of propellant that the case contains but gets away from the restrictions of an end-burning grain. Almost any desired thrust-time program can be obtained by using various geometric shapes for the cross section of the perforation. Grains can be designed for level thrust during the entire period of propellant burning, producing constantly increasing acceleration since fuel consumption lightens the rocket. In other programs, an increase in thrust with time produces very large increases in acceleration. Thrust can be made to decrease pro­ gressively, producing constant acceleration. Still other designs produce a two-level thrust program for booster and sustainer action from a single motor. In most cases, level or nearly level thrust programs are employed. To obtain this condition, the original burning area must be maintained throughout the period of operation. In a grain with a circular perforation and burning on both ends this condition can be met by selecting the proper length, grain diameter, and web thickness. As burning progresses, the radius and therefore the area of the perforation increases. But this area increase is compensated by the shortening of the grain as the ends burn. Though the thrust-time program is not completely level, the initial and final thrust can be made the same. An advantage of the circular perforated grain is that the residual propellant burns out at all points at the same time. Thus, all the pro­ pellant makes its full contribution to thrust. This is not true of designs to be discussed later. However, for level burning the required dimensions are such that only 73 per cent volumetric loading density is achieved. Other designs greatly exceed this value. Circular perforated grains have been used extensively in propellant development programs as ballistic test motors. Flight-type motors have also been designed with circular perfora­ tions. High loading density can be obtained, but special design features must be incorporated for level burning such as a head-end web or partial restriction of the ends of the cylinder. This type of grain is not used as frequently as those of other designs.

C. Star Perforations

More generally used are the grain designs in which the center perfora­ tion is a star instead of being circular. The star can be designed so that its perimeter stays relatively the same as propellant burns and therefore thrust remains level. Burning on the ends can be restricted by coating

SOLID PROPELLANT ROCKET

59

TECHNOLOGY

with plastic and then level burning is independent of grain length. It is more usual to allow burning on the nozzle end and cover the head end of the case with propellant, to a depth usually less than the full web thick­ ness. The increasing hemispherical surface of the head end propellant web compensates for the shortening of the grain as the aft end burns. Very high volumetric loading density can be achieved with star perforated grains. However, there are certain restrictions on the design. For a level thrust-time program, the perimeter of the star must equal the outside perimeter of the propellant grain. Because the flame does not reach the case at all points simultaneously, there is unburned propellant between the star points at web burnout (Fig. 3). This propellant is known as "sliver." Because of the great reduction in area at web burnout, the slivers burn at reduced pressure, resulting in a loss in efficiency. In grain design the "sliver loss" is kept to a minimum except for special reasons. Star design can be made in which the burning area approaches a circle shortly before reaching the

Port

FIG. 3. Configuration of internal burning grain (showing the burning pattern).

60

I I . W . R I T C H E Y AND J . M . MCDERMOTT

case. This minimizes sliver loss, usually at the expense of progressive increase in thrust for the last half of the burning time. D. Thrust Programs

Thrust ( l b )

Figure 4 shows various star designs and the thrust-time program each produces. Five- and six-point stars are used extensively. They can be

Thrust ( l b )

Burning time (sec)

Thrust ( l b )

Burning time (sec)

Thrust ( l b )

Burning time (sec)

FIG. 4. Grain configurations and thrust-time records.

made level burning, slightly regressive, or slightly progressive with head end propellant or a tapered perforation, larger at the nozzle end. As the number of star points is increased to seven or more, regressivity usually becomes greater. Figure 3 above shows a special case known as a double web grain. The propellant between the star points is twice as thick as the web and is entirely gone at web burnout. This grain produces high thrust from the large-burning-area—large-throat combination but also gives a sliver weight of 7 to 15 per cent of the propellant. Many other designs, such as double anchors and cog wheels, have been used for special purposes.

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E. Port-to-Throat Ratio

The center perforation is called the port. There is a differential pressure between the head end of the motor and the nozzle end, causing gas flow in the port. When the area of port is four times the area of the throat, gas velocity in the center perforation is safely low. However, at a port-to-throat ratio of 1, the velocity in the aft portion of the port is Mach 1. No conventional rocket operates at this ratio, the approximate lowest being 1.5 because of a phenomenon called "erosive burning." This phenomenon is caused by hot gases flowing at high velocity parallel to the surface of the burning propellant. Increased heat transfer to the surface occurs at the aft end of the perforation, greatly increasing burn­ ing rates in this region, increasing correspondingly the pressure and thrust. Equations of limited usefulness have been developed to describe erosive burning. The undesirable effects which result are a very high pressure peak at the beginning of burning and a premature regressive burning at the end. The high pressure necessitates a stronger, heavier case than otherwise needed and the change to regressive thrust is frequently un­ desirable. Designers sometimes enlarge the aft 10 per cent of the length of the perforation to increase the port-to-throat ratio and decrease or eliminate erosive burning. The port-to-throat ratio is one limiting factor on how much propellant can be put into an internal burning star design. Another limiting factor is the mechanical property of the propellant, especially in case-bonded grains. As mentioned before, the coefficients of thermal expansion for propellants and steels differ by a factor of approxi­ mately 10. Therefore, in cooling after removal from the curing oven, the grain is strained by thermal shrinkage. Being restricted from shrinkage on the outside by the bond to the case, the star perimeter must enlarge; but as can be seen by inspection of the figures showing grain designs, the strain will be nonuniform, concentrating in the star points. The strain induced must obviously be below the strain limit of the propellant. The propellant must be able to undergo this strain at a stress less than the mechanical strength of the propellant in tension. Otherwise, the grain will crack. The nearly straight-sided, square-ended star point with filleted corners is designed to distribute stress. Points with a sharp V-shape in­ crease the concentration. These last two statements are confirmed by photostress model studies. Either port-to-throat ratio or propellant mechanical properties may limit the thickness of the web. Rigid mathematical analysis of the re­ quired mechanical properties of star designs is extremely difficult. The treatment of the grain as a thick walled hollow cylinder has been done; it has been shown that there is a critical ratio between the diameter of

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the perforation and the outside diameter of the grain for any given set of propellant physical properties. If this ratio is exceeded, cracking will occur. This ratio is temperature dependent and is fixed by the lowest temperature to which the grain will be exposed either in transportation or storage. This ratio also imposes a limit on increase of burning time by the use of a thicker web.

F. Mass Ratio

The mass ratio is the weight of propellant divided by the total weight of the motor which includes propellant, case, insulation, nozzle, and igni­ tion system and which may include destruct system and auxiliary power unit for steering. Mass ratio is improved by high volumetric loading, high propellant density, and low inert parts weight. In the discussion of mass ratio, either mass or weight may be used, since mass = W/32.2 and the mass ratio is obviously the same as the weight ratio. High volumetric loading density is achieved by careful selection of star design, port-to-throat ratio and web thickness which is closely con­ nected with burning time. The density of propellants available today may 3 approach 1.8 gm/cm . Because of its importance, much effort has been expended in attempts to improve this figure. High propellant density, because of its relationship to both high specific impulse and high linear burning rate, provides the capability of high mass rates which contribute to high thrust-to-weight ratio. Low inert parts weight is achieved by making the case of alloys of very high tensile yield strength. The develop­ ment of these alloys has made important contributions to solid fuel rocket efficiency. Further improvements are in prospect. Nozzles, though usually quite massive, can be lightened by careful selection of materials and design. The use of plastics as structural materials is a result of their high strength-to-density ratio. Mass ratios of solid fuel motors ranging from 0.85 to 0.94 can be achieved with thrust-to-weight ratios about 4. However, optimization of all factors, pressure, case strength and thickness, burning time, thrust-toweight ratio and buckling or aerodynamic loads often dictate designing to the lower figure. This mass ratio is lower than that of turbopump liq­ uid motors when the system is fueled to provide a burning time in the range of 120 sec, resulting in a low thrust-to-weight ratio and attendant g losses when used in first stage applications. The equation for velocity at burnout emphasizes strongly the impor­ tance of both mass ratio and specific impulse. Under ideal gravity-free and drag-free conditions

where

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VB = /.pffclofrOTo/F,)

(11)

hp = specific impulse gc = 32.2 (standard Earth gravity) T^o = mass at launch in pounds WF = mass at burnout in pounds (Wo — WF = mass of propellant) If we add the terms to apply this ideal equation to a rocket in flight VB = 2.3O26J8P0C logio (WO/WF)

- jT 0 cos a dt -

£

dt

(12)

where a = angle between the trajectory and the vertical Wi = mass at any time t in pounds tB = burning time fdrag = drag force g = gravitational acceleration acting on the rocket at any instant t = time This points to the possibility of optimizing the burning time in order to optimize the system for losses due to gravity and drag, giving higher burnout velocities. IV. Ignition of Solid Propellant Rockets A. Pyrotechnic Igniters

Solid propellant rockets can be ignited by means of an electric squib and a pyrotechnic mixture. The squibs contain a very small amount of primary combustive material and a grain or so of pyrotechnic powder. This charge and the burnout wire are encased in a metal cup crimped tightly to prevent contamination of the charge by vapors from the pro­ pellant or by moisture. The squibs are designed not to fire until a certain critical electrical energy is applied. This allows continuity testing with­ out danger of premature ignition. I t also prevents the squib from firing from stray induced currents from electronic gear or power lines in the area. However, when the critical electrical energy is applied the time from closing the switch to firing is only a few milliseconds. The pyrotechnic mixture is made up of an inorganic oxidizer and a metal powder. At one time black powder was used extensively for ignition but its use has been largely discontinued. It contributed large quantities of hot gas with an undesirable brisance.

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To eliminate the undesirable features of black powder, the igniters are designed to deliver a cloud of hot particles but relatively little gas. The metal fuels that have been used are aluminum, magnesium, titanium or zirconium powders or finely divided boron. The most generally used oxidizers are potassium perchlorate or nitrate. A small amount of plastic is used to bind the mixture together. Burning is very rapid. Smooth igni­ tion is achieved in less than 50 msec. The design of the igniter depends on the design of the motor and its mission. It must be designed and positioned to ignite a large portion of the burning area at the same time. This results in a smooth rate of pressure rise. However, it must not be too large for its mission. The igniter causes a pressure rise in the motor from its own heat and gas and by heating the air in the case. This is desirable since solid fuels ignite better and burn faster under pressure. But there must not be an overpressurization which might rupture the case or damage the grain. For certain missions debris from the igniter, ejected from the nozzle, can cause damage to structures or injury to personnel. Therefore, it is desirable to have all or as much as possible of the igniter structure burn up in the motor chamber. Cloth or plastic bags containing the squib and pyrotechnic mixture in pellet form have been used for small rockets. For large rockets where the grain perforation is long, a perforated plastic tube may be used to contain the pyrotechnic. This may be half or more of the length of the perfora­ tion. Another design employs cloth or plastic covered on one side with pyrotechnic paste and rolled into a tube. The perforated tube and the rolled tube have to be supported by the star points to avoid damage from jars or vibrations in motor handling. Whatever the design, two squibs or more are used to insure reliability and uniform ignition throughout the mass of pyrotechnic. Pyrotechnic igniters are usually inserted through the nozzle and the wires to the squibs are led out the nozzle. However, head end positioned igniters are common. B. Pyrogens

Another method of ignition employs a small rocket motor to ignite the main charge. The small motor is called a pyrogen. Use of a pyrogen gives smooth hot gas ignition without brisance. The exhaust from the pyrogen is directed into the center perforation, usually from the forward end. Fast burning propellants are used at moderately high pressures to obtain a high mass discharge rate. The exhaust from the pyrogen ignites the main grain smoothly and reproducibly. The design used for pyrogens is, in general, similar to the main charge.

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Very large surface areas are used with thin webs since the pyrogen only has to operate for 100-300 msec. The DeLaval nozzle and its expanding exit section may be eliminated since sonic flow of the exhaust gases is sufficient. The pyrogen is usually inserted through a boss in the forward dome and attached by a flange on its own head end. It must be sufficiently large to ignite the main charge smoothly, rapidly and reproducibly but not large enough to cause overpressurization. Every effort is made to make the pyrogen case and nozzle as light as possible to decrease the weight penalty. For very large motors the use of a pyrogen provides a better method of ignition. A pyrotechnic igniter, though consumed or ejected upon ignition, would have to be very large and heavy, presenting prob­ lems of support during transportation. C. Ignition in Space

It has been mentioned before that solid fuels burn better and faster under pressure. In most cases, they do not burn at all at pressures of 100 mm Hg or less. This, at one time, appeared to be a potential problem for space applications. However, closures have been designed to retain igniter pressure till it can build up to the point at which the grain ignites. Squibs perform reliably at very high vacuum. Use of solid fuel retrorockets to return vehicles from orbit have demonstrated the reliability of ig­ nition of solid fuels in space. V. Testing of Solid Propellant Rockets A. Static Tests

Static firing of solid fuel motors is performed extensively during all development programs to test propellants, inert components, and design concepts. Post development static tests are used to demonstrate quality assurance. Also, static tests are conducted for quality control of the pro­ pellant. Solid fuel motors are tested in horizontal or vertical stands. Vertical testing is usually performed with the head end down, requiring very sim­ ple test stand construction. Multicomponent stands are used to measure side forces developed by thrust vector control devices. The test cells are built of reinforced concrete. Instrumentation and personnel are in an underground or well barricaded block house. The minimum measurements during static test are thrust, pressure at one or more points in the chamber, and burning time. These values define the ballistic performance. In addition, depending on the purpose and im-

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portance of the test, other measurements may be taken. Thermocouples may be attached to the case and nozzle, strain gages may be used on critical areas of the case, and photographic coverage from various angles may be provided. B. Environmental Tests

The kind and extent of environmental tests are governed by the envi­ ronment in which the motor will have to perform. Storage for periods up to months at very low temperature without cracking or bond failure may be a requirement of military rocket motors. Also, storage at tempera­ tures up to 140°F without a loss in ballistic performance is required for service in the tropics or deserts. For many types of motors salt spray tests are required. At some time during motor development the propellant and rubber or plastic parts may be exposed to various strains of fungus. For very large motors many of these tests are dispensed with because con­ trolled storage environment obviates the need. However, the ability to endure the variety of severe environmental conditions attests to the ruggedness of the solid propellant motors. C. Storage Tests

Since one of the most important capabilities of solid fuel motors con­ sists of being ready for immediate use after long period in storage, the aging properties of the propellant and any inert parts subject to deterio­ ration are carefully observed. The propellant is required to age without change in mechanical and ballistic properties. The inert components must not change in any way that would affect their performance. Aging investigations may be conducted at either elevated or ambient temperature. Motors and propellant samples are usually stored together. The motors are inspected and static tested after different periods of aging. The mechanical properties of the propellant are measured. Artifi­ cially high temperature storage has the advantage of providing aging data more rapidly. However, changes may occur under the influence of high temperature that would not occur for years or not at all at ambient. Since few motors will be stored continuously at high temperatures, the real proof of aging properties comes from ambient storage. D. Drop Tests

For small motors and even for some of several hundred pounds, drop tests are frequently specified. The drop may vary from a few feet up to

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thirty feet onto a concrete pad. The specification may require that the motor not ignite or explode. E. Detonation Tests

Solid propellants are frequently referred to as explosive materials. They are, not however, designed to be explosives in the sense that T N T or tetryl are explosives. Workers in the field of explosives distinguish between deflagration and detonation. Burning or deflagration is a process in which a surface decomposition once initiated generates gas and heat. Heat is fed back to the surface keeping the decomposition going until all the material is consumed, usually at a rate measured in inches per second. High explosives undergo decomposition which manifests itself as a high velocity shock wave known as a detonation wave. The velocity of this shock wave varies with the material and its density and physical condi­ tion from 2000 to 20,000 ft/sec. It should be noted that some materials, e.g., TNT, will burn smoothly with a sooty flame. But in a large mass, burning raises the temperature of the material faster than heat can es­ cape. The hotter it gets the faster it burns until enough energy is being generated to initiate a detonation. This is known as the deflagration-todetonation transition. The importance of knowing how a propellant will react under acci­ dental conditions has led to a variety of explosive tests. In the development of a new propellant, a sample confined in a 4 x 48-in. steel pipe is initi­ ated with a dynamite cap and 0.25 lb of tetryl. Several sections of steel pipe may be laid end to end or separated by either an air gap or sheets of cardboard. The explosive charge is set off at one end and observation is made of the number of sections through which the disturbance propa­ gates. There are various methods of determining whether a material det­ onates or merely burns rapidly. One way is to attach one end of a length of prima-cord several feet long at intervals along the length of the test samples or motor. A lead block is attached to the far end of the primacord. If detonation occurs, the wave is conducted to the lead block which serves as a witness to the fact of a detonation. However, much of the in­ terpretation depends on the judgment of an experienced operator, based on the amount of fragmentation and the size of the crater produced in the ground. In other tests, motors are placed in the middle of piles of scrap lum­ ber or pits containing gasoline which are then ignited. Also, motors are tested by placing a substantial charge of high explosive against the case and setting it off. Most composite propellants do not detonate. The ammonium perchlo-

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rate itself is an explosive, especially when mixed in the right proportion with an easily oxidized material. But the rubbery binder is nonexplosive and appears to damp out the detonation wave initiated by the high ex­ plosive. Also, the composites burn rather slowly, mostly below 1.0 in./sec and do not undergo the deflagration-detonation transition. This is fortunate for very large vehicles where a mass detonation would present safety problems. By the nature of their components, nitrocellulose and glycerol trini­ trate, double base propellants are more susceptible to detonation than the composites. However, the energy required to initiate a detonation is very large and these propellants are quite safe under ordinary conditions for handling and for use in motors. Pressure bursts, though frequently quite violent, are much less destructive than a detonation. VI. Solid Propellant Processing A. Extruded Double Base Propellant

Solventless double base propellants are made by adding glycerol trini­ trate to an agitated water slurry of nitrocellulose. The glycerol trinitrate is absorbed by the nitrocellulose. The other ingredients, stabilizers, and ballistic modifiers are added to the slurry. The water is then removed by centrifuging, the mixture is aged for several hours to allow complete dif­ fusion of the glycerol trinitrate into the nitrocellulose and then further dried with heat. It is treated on a differential roll mill followed by treat­ ment on even speed rolls, finally emerging as rolls of material 0.1 in. thick, 4.5 in. wide and many feet long. Preparation of the propellant grains is done in a hydraulically oper­ ated extrusion press. Both vertical and horizontal presses have been used. Rolls of solventless powder are stacked in the press after being preheated to 120°F. In vertical presses granular powder may be used. Both the ex­ truder barrel and dies are heated. The cylinder is evacuated to remove air and extrusion is performed at about 3000 psi. Large presses have a diameter of 15 in. and handle about 130 lb of propellant per charge. The size of the die is about one-third the diameter of the press. This limits the size grain that may be produced to 5-6 in. As the propellant emerges from the die it is cut to length and annealed to remove stresses. Washers of a plastic such as ethyl cellulose are ce­ mented to the ends to restrict burning. In some cases the outer diameter is sufficiently uniform but in other cases the grain must be machined to constant dimension. Final inhibiting is accomplished by wrapping in a spiral pattern with plastic tape which is cemented in place. With the

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burning area thus restricted to the central perforation the grain is ready to be inserted into its rocket case. B. Cast Double Base Propellant

The casting process for the manufacture of double base propellants is used in the manufacture of grains larger than those conveniently made by extrusion. The basic materials again are nitrocellulose and glycerol trinitrate. The nitrocellulose may be granules made by the solvent proc­ ess or small round particles made by the ball process. These types are manufactured extensively for small arms ammunition. In both processes the stabilizers and ballistic modifiers are added during manufacture of the powder. The first step in the preparation of the rocket fuel is the formation of a plastic mold into which propellant will be cast. Cellulose acetate or ethyl cellulose sheet is wrapped around a cylindrical mandrel of the proper size. The plastic sheets are cemented together to various thick­ nesses, possibly up to % in.? depending on the size and burning time of the grain. One end of the tube is closed with an equal thickness of plastic. The form, called a beaker, is then dried to remove residual solvent. The beaker is slipped inside a heavy walled metal container. The core or mandrel, usually of aluminum, for forming the central perforation is coated with a suitable release agent to prevent it from sticking to the grain. It is then installed in the proper position within the beaker, sup­ ported by attachments to the metal container. The granular or ball powder is poured into the beaker. Vibration may be used to insure uni­ form close packing. The container is then closed with a metal cover and sealed and the assembly is evacuated. This removes residual moisture and air either of which, trapped among the granules, could cause bubbles or voids. The glycerol trinitrate is then introduced through suitable openings in the bottom of the container and percolated through the entire mass of powder, an operation requiring considerable care. Too rapid addition disturbs the bed of granules; still, the rate must be rapid enough to allow introduction of the entire amount before swelling of the granules shuts off the flow en­ tirely. At the end of casting, the grain is allowed to cure at slightly elevated temperatures, 120-135°F, for periods up to several days. The granules swell from absorption of the glycerol trinitrate and coalesce, and the mix­ ture is thus converted to a single uniform grain. After the cure is com­ plete the grain is cooled gradually, the mandrel is withdrawn, and the beaker containing the grain is removed from the metal container. Plastic

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H. W. RITCHEY AND J . M. MCDERMOTT

inhibitor is applied to the free end. The grain is then inserted into the rocket case and is ready for firing. Other organic nitrates and nitro compounds are also used as additives in double base propellants. Some of these are guanidine nitrate, ammo­ nium picrate, and dinitro toluene. Diethylene glycol dinitrate has also found use as a plasticizer. Proper compounding of ingredients has altered the characteristics of double base propellants from rigid plastics to elastic materials suitable for case bonded grains. Ammonium perchlorate and metal powders are also being incorporated to improve oxygen balance and increase the heat of combustion. C. Composite Rubber Base Propellant

The manufacturing operation with composite rubber base propellants is very different from that for solventless or cast double base. The proc­ ess starts with case preparation. The inside of the case must be scrupu­ lously clean to insure a good bond. This is accomplished by grit blasting or wire brushing followed by vapor degreasing or solvent degreasing with trichloroethylene. Any required internal insulation is then bonded in place. Most case bonded grains have a so-called liner over the entire inside surface of the case. The liner is a polymeric material, often the same polymer as the fuel binder, applied to a thickness of 0.04-0.1 in. The liner is loaded with inert fillers such as carbon black or finely divided silica. It serves as an adhesive, and protects the case from the hot gases during sliver burnout. For small motors the liner may be applied by a pour-in pour-out method but it is usually sprayed on like a paint. The propellant may be applied to a fully cured, partially cured, or uncured liner, depending on the compound. If cure or partial cure is required it is accomplished in circulating hot air ovens. The oxidizer, if ammonium perchlorate, is usually very dry as re­ ceived and is ready for grinding without further treatment except a pre­ cautionary screening for foreign objects. Grinding has been performed in various types of hammer mills. This operation is performed behind a barricade for protection of the operators. The fineness of the ground ma­ terial depends on the particle size of the feed and the speed of the mill and is quite reproducible. Most propellants use an oxidizer consisting of a mixture of two or more particle sizes to obtain better flow properties and for burning rate control. The different sizes may be blended after grinding or added to the mixer separately. The proper amount of the blend, or the separate sizes, is carefully weighed out into drums or other forms of transportable containers and transported to the mixing area.

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The polymer for the fuel binder is carefully weighed into the bowl of a pre-mix mixer. This is usually a two paddle epicycloid mixer such as those used in bakeries. The other ingredients, burning rate catalysts, metal powder, plasticizers and, in some cases, surfactants are weighed and added to the polymer. The curing agents may be added at this point or they may be withheld until the final stages of mixing. Since the pro­ pellant must retain flow properties for 4-6 hr after mixing, the rapid onset of cure cannot be allowed. The pre-mix is stirred for a specified time and is ready for the propellant mixing operation. Propellant mixing is performed in Baker-Perkins mixers of large size. The sigma blade type has found most use, though other types of blades are also employed. The oxidizer may be added to the mixer and the premix run in on top or the pre-mix may be added first and the oxidizer brought in through a chute from an overhead bin. The mixing operation is performed behind a barricade with no personnel present. Mixing time is specified for each type of propellant. At a predetermined time in the mixing cycle steam or hot water is turned into the mixer jacket so that the propellant will reach the proper temperature, 130-150°F, at the end of the specified mix time. The elevated temperature is required to improve the fluidity of the mix. The mix may then be dumped into a large funnel whose bottom out­ let is a flat plate, perforated with slits 0.125-0.25 in. wide. The funnel is attached to the cover of a large jacketed can called a casting can. The can is positioned in a pit in front of the mixer and the propellant pours directly from mixer to funnel. The casting can is evacuated, and as the propellant flows through the slits in narrow ribbons the entrapped air is removed. At the end of the operation, the deaeration head is replaced by a pressure head and the propellant is ready for casting. Hot water circu­ lates through the jacket of the casting can to maintain propellant tem­ perature to the end of casting. The lined motor is fitted with an extension sleeve at the aft end and the mandrel for the central perforation is put in position. The mandrel and sleeve are covered with mold release to prevent sticking to the pro­ pellant. Large motors are cast in pits below ground but small motors are filled at floor level. A hose is fitted to a diaphragm valve in the bottom of the casting can and led into the motor. The propellant is forced through the hose by air pressure and the end of the hose is raised continuously as the propellant surface rises. For small motors, dozens may be filled from the load of one casting can. For large motors, the casting can may be changed several times and one batch of propellant cast on top of the previous batch. The motor is filled until the propellant rises into the extension sleeve.

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Either pit curing or oven curing is used to convert it to a uniform rubbery material adhering tightly to the case wall. The ovens or pits use circulat­ ing hot air, the time and temperature and the rate of raising and lowering the temperature depending on the type of propellant. Temperatures may vary from 135-185°F and time from a few hours to a few days. Tempera­ tures of 300°F have been used for some types of propellant. After cure the motor may be removed from the oven at once or go through a programmed cool down. The mandrel may be withdrawn after cooling or with the propellant still hot. When completely cooled, the ex­ tension sleeve is removed and the grain cut back by a remotely controlled cut-back machine. This operation is necessary to effect precise control of grain length or weight. When the aft closure, nozzle and ignition device have been installed, the motor is ready for service. VII. Inspection and Quality Control A. Raw Materials

The assurance of reliability and reproducibility of performance of solid fuel rockets starts with the acceptance tests of the raw materials. The polymers for the fuel binder of composite propellants, the adhesives and sealants are all procured according to tight specifications on both chemical and physical properties. The common chemical tests include moisture, ash, impurities known to be detrimental, and number of reac­ tive groups per unit volume or weight. The physical tests include viscos­ ity and density. Frequently, acceptance may also depend on the results of tests of cure time and of the tensile strength and elongation of the cured polymer. The explosive ingredients for double base propellants undergo several in-process inspections and tests and final acceptance depends on chemi­ cal analysis, particle size, and stability. Some solventless powder is also accepted on the tensile strength of the rolled sheet. Oxidizer is accepted on the basis of chemical analysis and particle size measurements. Chemical analysis is used to determine moisture, ash, active ingredients, and impurities that may sensitize the oxidizer, such as chlorate and bromate, and undesirable cations of the alkaline earth group. Particle size is usually determined with standard sieves. The re­ producibility of the method is good, but the obtaining of a uniform sam­ ple from large lots of oxidizer presents a fair amount of difficulty. The ingredients or materials going into other inert parts are also care­ fully tested for composition or functional qualities.

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B. In-Process Controls

The number of in-process controls used for composite propellants is not large. Correct weighing of ingredients and adherence to standard op­ erating procedures are important and usually are certified by process inspectors. Viscosity is measured at the end of mixing. Total solids may also be determined at this point. Uncured strands of propellant in paper or plastic straws may be burned after the deaeration step to prove con­ formance to burning rate requirements. For large motors a number of small test motors are frequently cast from each batch for quality assur­ ance tests. This is not ordinarily done for small motors where a statistical sample from each lot can be test fired. It is usual to cast from each pro­ pellant batch one or more bulk samples to be cured with the motor or lot of motors for tests of mechanical and physical properties. Molded reinforced plastic parts are inspected for conformity to di­ mensions specified on the drawing. An X-ray inspection is frequently made for voids or other defects. Such parts may be destructively tested for resin content, residual volatiles, tensile strength and Young's modulus. In-process inspection of the case starts with chemical analysis and ultrasonic inspection of the billet. A sample of the billet may also be taken to determine heat treat response. Inspection is again performed after forging or rolling. After these operations the case may be formed by deep drawing, spinning, or flow turning. At the completion of these operations the trimmings are saved for coupons that go through heat treat with the case. The welds are inspected by X-ray, magnetic particle, or fluorescent penetrant. After final heat treat and tempering, the case is inspected by magnetic particle or fluorescent penetrant and the heat treat coupons are tensile-tested. Finally, the case and closure are inspected for dimensional adherence to specification and hydraulically proof-tested to the 3-sigma expected maximum internal pressure plus a safety factor. A final magnetic particle or fluorescent penetrant inspection may be re­ quired for large cases to find any cracks initiated under strain in proof test. C. Cured Propellant

Most of the testing and inspection following the cure of propellants are for physical and mechanical properties and for the integrity of the grain. The polymers used in composite propellants, once cured, resist re­ solution for analysis by wet chemical methods. Being highly loaded with inorganic materials, they are poor subjects for methods depending on

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light transmission. No chemical changes occur during the cure of double base propellants. Using the bulk samples cured with the motors, it is usual to determine the tensile strength and elongation of the propellant and the modulus (or compliance) up to the limit of the elastic response of the material. Being visco-elastic materials, Young's modulus does not apply as it does for truly elastic bodies. Density is also determined. Small motors are weighed and a certain number are fired as a check on burning rate, Kn and specific impulse. For large motors, the small motor cast from each batch are static tested for proof of quality. D. Radiographic Inspection

The great importance of having the propellant grain uniform and free of flaws make radiographic inspection a necessity. Various methods have been employed depending on the size of the motor. A cobalt-60 source, standard X-ray machines, Van de Graaff machines, fluoroscopes, and linear accelerators have all been used. The film may be placed in the cavity or the exterior of the case. Bubbles, cracks, and low density areas are detected if they exceed a certain minimum size. Separations between the propellant and liner or between liner and case are difficult to detect. This requires a tangential exposure and the pictures are frequently difficult to interpret. Fortunately, this is an un­ common type of defect and not ordinarily a major cause of concern. VIII. Grain Defects

There are three types of grain defects that may be serious: (1) low density areas, (2) cracks, and (3) bubbles of entrapped air. Low density areas, if they are spongy from trapped air or gas, burn at a vastly in­ creased rate. This can result in exposure of the case wall to the flame for a much longer period of time than the liner was designed to withstand, bringing about case failure from the internal pressure. The effect of a crack is to increase the burning area and therefore the pressure, possibly causing the motor case to burst. The effect of a bubble is the same as that of a crack. Fortunately, good design and carefully controlled manufac­ turing methods have reduced the incidence of cracks and bubbles to tol­ erable limits. The size defect that can be tolerated depends on the size of the motor and the location of the defect. It is usual to specify the largest acceptable defect size for each motor, depending upon the tolerable in­ crease in surface, and more especially upon the tolerable pre-exposure of the case wall to hot gas before burnout.

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IX. Thrust Vector Control and Thrust Termination

There are several ways of controlling the direction and attitude of a solid fueled rocket. Moderate forces may be required to correct for me­ chanical malalignment or to compensate for gusts at lift off. After igni­ tion, solid fuel motors attain full thrust in milliseconds so that the period when this control is required is very short. At altitude, pitch and roll con­ trol is required as well as steering. In most cases the forces required for thrust vector control (TVC) are not large and can be generated by sim­ ple devices (Fig. 5). One of the oldest methods of TVC is the use of four jet vanes. Jet vanes, when applied to a single nozzle, have the advantage of providing roll control torques as well as side vectors. This results from deflection of the exhaust as the vane swings on a pivot on the edge of the exit cone. There are two principal disadvantages to the use of jet vanes. Large motors require vanes with large surfaces and large cross-sections, result­ ing in high drag losses in the exhaust jet. Since the vanes are constantly immersed in the high temperature, high velocity gas stream, they are eroded rapidly, introducing a materials problem. The materials problem is partially avoided by the use of the jetavator. This device is the central zone of a sphere, usually mounted on gimbals. It dips into the exhaust jet in the direction required to provide the nec­ essary thrust vector. In contrast to the jet vane, the jetavator is immersed in the exhaust only for the time the control forces are needed. It has the disadvantage of inducing relatively high drag loss during the period that it is in use. Also, it is not capable of providing roll control unless the rocket motor is fitted with multiple nozzles. A third TVC method uses a hinged nozzle in much the same way a liquid motor uses a gimbaled combustion chamber. In the case of a solid motor, the nozzle is connected by a flexible coupling or hinged joint to the combustion chamber. I t is mounted on gimbals so that the exhaust jet can be diverted in any direction. This device is the most efficient of the three for cutting drag loss. The mechanical problem of providing seals against the hot, high pressure gas is severe. Also, roll control can only be pro­ vided if the motor is equipped with multiple nozzles. A fourth method is the injection of a fluid (gas or liquid) into the ex­ haust stream at a point aft of the nozzle throat. The fluid is introduced through small openings in the body of the expansion cone. With a num­ ber of openings, evenly spaced around the cone, to be opened and closed as needed, differential pressures would be generated at the wall to provide side thrust. It is necessary to carry along a reservoir of gas or liquid injectant, or to bleed gas from the main chamber to avoid the weight pen-

FIG. 5. Thrust vector control devices.

Jetevotor

92, xxoHHaaoH *K T QNV ASHOUH *AV *H

SOLID PROPELLANT ROCKET TECHNOLOGY

77

alty. The combustion bleed system presents severe materials problems in ducts, manifolds and valves to contain and conduct the very hot, high pressure gas. The best method for TVC can only be selected after optimi­ zation studies reveal the trade off between flexible nozzles or fluid injec­ tion. Thrust termination can be provided by generating an exhaust jet opposite to the direction of the main propulsion stream. Under these conditions, a reverse thrust can be generated that is equal to or slightly larger than the main forward thrust. Thrust termination devices are shown in Fig. 6. The reverse thrust is accomplished by opening reversal ducts connecting into a plenum chamber at the aft end of the engine. Forward end ducts may also be used, but usually must be placed at an angle to the main axis so that the pay load is not subjected to the heat of the reverse jet. Both aft-end and head-end reversal devices require an arrangement for rapid opening of the activation ports. Failure to obtain rapid and simultaneous opening of several ports can generate unbalanced side forces sufficient to tumble the missile. The ports are opened by bursting a diaphragm with a squib-actuated explosive charge. The addi­ tion of the plenum chamber and its reversal nozzles adds substantially to the weight of the motor. A third method of thrust termination involves quenching of the pro­ pellant burning. This is done by opening a new nozzle area A2, which is larger than the initial throat area. This, in turn, is accomplished by re­ leasing the nozzle by an explosively activated device. A shock expansion wave is set up inside the combustion chamber that extinguishes the pro­ pellant. The expansion wave cools the intermediate mixing zone, lowering the burning surface temperature below the pyrolysis temperature, and thus terminating the flow of gas into the flame front. Under atmospheric conditions the propellant will frequently reignite because of heat in the inert parts. At high altitudes reignition does not occur because combus­ tion cannot resume at very low pressures. This extinction process has the advantage of incurring little or no weight penalty. Also, gas flow is uniform along the main thrust axis so that overturning moments are minimized. The disadvantage is that it causes a transient surge of thrust that could, in some cases, be damaging to the payload. This surge is caused by the larger port area during the transient decrease from combustion chamber pressure. Thrust increases momentarily to a value roughly equal to F1{A2/A1). If it is assumed that the expansion wave travels forward at a rate equal to the speed of sound in the combustion products, around 3000 ft/sec, the duration of the transient can be approximated by dividing this value into the length of the chamber. After the peak, pressure decays rapidly to ambient.

FIG. 6. Thrust termination devices.

Blow- Off* Nozzle

xxoHHaaoH *K T QNV ASHOUH *AV *H

SOLID PROPELLANT ROCKET TECHNOLOGY

79

Once any of these methods of thrust reversal or termination have been employed the motor is no longer useful since it cannot be restarted. This is one of the disadvantages that has not been overcome in solid fuel rockets. Also thrust modulation, though possible, presents extremely se­ vere materials problems. It could be accomplished with auxiliary jets whose area could be varied as required by means of pintles or irises. X. Solid Fuel Rockets for Space Missions

Solid fuel motors have been used in space applications as upper stages of such space carrier vehicles as Vanguard and Juno 2 and as spin rock­ ets and retrorockets. They have performed reliably in these missions. It appears that their characteristics are especially applicable to first stage requirements of large space carrier vehicles. Not all the factors important for first stages apply to the same degree to upper stages. The upper stage motors do not require high thrust-toweight ratio. They work in an area of zero gravity or at least greatly reduced gravity. Thus the highest mass ratio consistent with over-all re­ quirements can be used. In the atmosphere the thrust coefficient increases with the pressure drop from chamber to ambient pressure at the exit. In space this pressure drop becomes infinite. Therefore, the motor can be operated at low pressure and still attain high specific impulse. This al­ lows weight savings in the case since it no longer must contain high pressures. It also allows longer burning times where they are needed. Thrust in space can also be increased by lengthening the nozzle expand­ ing section to increase the expansion ratio. This is effective until the ad­ ditional weight becomes a penalty. Though the mass ratio and thrust of solid fuel upper stages can be im­ proved by these means, turbopump liquid fuel motors have advantages for these applications. They may use cryogenic propellants of very high specific impulse. They can be fueled for long burning times, giving high mass fractions. They are capable of being stopped and restarted and of thrust magnitude variation. However, starting or restarting in a zero gravity field presents problems unless the free surface of the fuel is con­ trolled. Solid fuel motors, on the other hand, are structurally stronger. They are much less complicated, resulting in high reliability. No starting dif­ ficulty has been experienced in a zero gravity environment. If optimized for mass ratio, high thrust, and proper burning time, solid fuel motors have their own advantages for upper stages. Carrier vehicles contemplated for space missions may be up to 100 ft long, 20-25 ft in diameter, and may weigh 2,000,000 lb. Larger motors

80

H . W . R I T C H E Y AND J .

M.

MCDERMOTT

could be built for heavier payloads. Parametric studies have shown that the first stage of optimum design will impart a velocity of about 5000 ft/sec to the payload, though larger increments could be obtained at some sacrifice in weight (Fig. 7). The payload in this case will be all upper 1——

,

,

4

6

_ —

Launch Weight + Payload Weight

40.

0

2

Velocity Change - I 0

8 3

10

12

ft/sec

FIG. 7. Velocity change of first stage vs the ratio, launch weight to payload weight.

stages, interstage structure, and manned capsule or instrument package. The ratio of launch weight to final payload package weight will be 20 to 40 for a two-stage vehicle for a 300 n. mile orbit. For a three-stage vehi­ cle in which the first and second stages contain solid propellant and the third stage liquid oxygen and liquid hydrogen, it will be 20 to 30. For a three-stage vehicle to achieve escape velocity, ratios above 50 are neces­ sary for any known combination of existing chemical propulsion systems. The optimization of space carrier vehicles is an extremely complicated operation because of the many factors which must be reconciled. Clas­ sically, the most efficient performance is obtained with high thrust-toweight ratios and short burning times. But gravity and drag impose restrictions on these values. Manned flight imposes limitations on accel­ eration. It is generally assumed that the structural integrity of the missile and weight penalties involved in stronger construction limit the dynamic 2 pressure to a maximum of about 1200 lb/ft . Then a first stage thrustto-weight ratio of approximately 1.8 is the maximum allowable and a first-stage burning time of over 70 sec becomes desirable (Fig. 8). If 1200

81

Payload to Launch Weight Ratio (lb/lb)

SOLID PROPELLANT ROCKET TECHNOLOGY

3

Burn-out Velocity ( I 0 ft/sec)

FIG. 8. Parameters for first stage solid fuel space carrier vehicle. 2

lb/ft is considered excessive, a lower F/W ratio must be established and burning time must increase (Fig. 9). According to these data, a firststage with F/W = 1.8 and a burning time of 70 sec, weighing 2,000,000 lb would produce 3,600,000 lb of thrust. The capability of solid fuel rock­ ets to produce thrusts of this magnitude should now be examined.

Payload to Launch Weight Ratio (lb/lb)

First Stage F/W Ratio

0.030

0 025

1

10

1

20

30

1

40

1

1

50

60

1

1

70

80

1

90

1

100

Burning Time (sec)

FIG. 9. Optimum burning time for first stage space carrier vehicle.

1

110

1

120

82

H. W. RITCHEY AND J . M. MCDERMOTT

Designing a space carrier vehicle is not a simple case of scaling up an existing motor. A useful example occurs in the scale-up of a smaller mo­ tor by a factor S from size 1 to size 2. The length, diameter, and burning time increase S times = L/2 SLi D2 = SD! (13) tm = Stsi 2 Because thrust depends on the area of the nozzle, it increases by S 2

F2 = FXS

(14)

Weight depends on volume and increases by S W2 = TfiS

3

3

(15)

Then F/W = 1/8 (Ft/Wi) (16) The XM-19 Recruit motor was used as the second and third stages of the Lockheed X-17. It was designed to gather aerodynamic data at ex­ tremely high velocity in dense atmosphere. The requirement of high thrust and small diameter was imposed to overcome drag forces. It is not an ideal motor to scale up but because of its high thrust-to-weight ratio will serve as a good illustration. Three of the parameters of the XM-19 are: Diameter, in. Weight, lb Thrust, lb

9 350 35,000

To illustrate the effect of scale factors on thrust-to-weight ratio, this motor will be scaled up by a factor of 50. The new diameter would be 2 37.5 ft. The thrust would now be 35,000 X 50 = 87,500,000 lb, and the stage would weigh 43,750,00 lb, resulting in a F/W of 2. Such a monster would lift a truly enormous payload. There is no known theoretical rea­ son why such a motor could not be built, but motors of this weight are not being considered at present. To obtain 3,600,000 lb thrust mentioned above the scale factor would be approximately 10.5. However, F/W would now be 9.5, producing higher acceleration than man could stand. This indicates some of the features that must be optimized for big solidpowered stages. Length, diameter, and mass fraction must be reconciled to produce the desired thrust, F/W ratio, and burning time. Another way to construct a large space carrier vehicle is to bundle a number of motors. There will be some decrease in F/W ratio over that of the individual motors because of the weight of the structure required to hold the bundle together. The parameters for the motors in the bundle

SOLID PROPELLANT ROCKET TECHNOLOGY

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would have to be optimized as carefully as for a single large vehicle. Study of all the propulsion requirements, including costs and handling methods, would dictate the choice of a single large motor or a bundle of smaller motors. The capability of both staging and bundling solid fuel motors has been demonstrated, most vividly in the Lockheed X-17. The first stage was an X M - 2 0 motor, the second stage a bundle of three X M - 1 9 (Re­ cruit) motors, and the third stage was a single XM-19. Propulsion for this missile performed reliably in all flights, totaling more than 3 0 . Stage separation occurred as programmed and ignition of the three secondstage motors occurred simultaneously, as did burnout, resulting in no deviation from the planned course. As mentioned before, there are no known theoretical limits on the size of a motor that can be manufactured. Practical limits arise from the F/W ratio which must be high enough to launch the vehicle and boost it to the velocity required for the mission. Other limits on both size and weight come from requirements for transportation and handling. Rocket motors up to 8 ft in diameter weighing 50,000 lb may be trans­ ported on highways without restriction. For motors up to 1 2 ft in diame­ ter and 4 0 0 , 0 0 0 lb, special multiaxle trailers and special routing would be required. Present railway facilities can handle motors 1 0 ft 8 in. in diameter weighing 1 5 0 , 0 0 0 lb. Special equipment and construction would enable the railroads to carry motors up to 1 4 ft in diameter and weighing 5 0 0 , 0 0 0 lb. Sizes and weights greater than this apparently cannot be han­ dled by railroad. Water transportation raises the limitation on size. The Saturn 1 is presently being moved by water from Huntsville, Alabama to Cape Ca­ naveral. However, flotation techniques will be required for handling and erection of the solid fuel motor at the launch site because of the great weight. Water transportation does restrict the location of both manufac­ turing plants and launching sites. Also, the motor can be made in segments light enough to transport or it can be assembled as an integral unit at the launching site. To manu­ facture large motors, batches of propellant are mixed and poured into the case on top of each other much as large concrete dams are poured. This multibatch casting has been practiced in large motor manufacturing for many years. The propellant cures to a monolithic grain with no seams or boundaries between batches. Large motors are usually manufactured in pits in the ground and curing is done without moving them. Loading very large motors at the launch site presupposes that methods and mate­ rials are available for case manufacture at the same site. Certain alloy steels are good candidates but some methods development may be re-

84

H. W. RITCHEY AND J . M. MCDERMOTT.

quired. Also, the use of resin-bonded glass fibers appears to be feasible. Mobile propellant mixing facilities, carried in trailer trucks, present no problems, so the site could be located anywhere that proves desirable. The segmented motor is manufactured in smaller, more conveniently handled segments which are assembled at the launching site. The seg­ ments can be made as large and heavy as can be conveniently handled. They would be loaded and cured in the usual way. This system would make use of existing propellant plants rather than the mobile plants for loading at the launch site. Case fabrication could be done at existing metal-working plants. The launching sites could be located anywhere that transportation and handling facilities permit. Operating flexibility is gained with segmented motors because the number of longitudinal seg­ ments can be increased or decreased to provide the thrust required for a particular mission. However, the design of joints between the segments presents problems of strength, rigidity, and gas tightness. There is also some weight penalty and some increased cost brought about by the use of joints. Small, solid fuel motors are used for space applications that require action within milliseconds after the switch is closed. The escape system that propels the manned capsule away from the vehicle in case of a mal­ function is powered by solid fuel rockets. Solid fuel retrorockets are used to bring a man or instrument package back from orbit. They are used as spin motors for large missiles where required. Also, stage separa­ tion is accomplished by the use of solid fuel motors. An area in which solid fuel rockets can be expected to become in­ creasingly important is in the return from the Moon or planets. The motors are very rugged and withstand severe jars, vibration, and expo­ sure to radiation up to a high limit without damage to case or propellant. The propellants are quite insensitive to shock. Their use might well im­ prove the capability to survive a lunar landing. Protection from the lunar temperatures would be required, as it would for men and instruments. In a reasonable temperature environment, the propellant does not deterio­ rate in storage. It is nonvolatile and noncorrosive. The simplicity, reli­ ability, and instant readiness of the solid fuel motors would be most im­ portant in the propulsion system designed for the return to Earth. One extremely important advantage possessed by solid fuel rockets is that they are ready for launch at the exact instant that the Moon or planets are in their most advantageous positions. This instant readiness results from the simplicity of the system. The countdown time is reduced to that needed for a few simple checks of the electrical ignition circuit and the thrust vector control system. Solid fuel rockets can be fired a few

SOLID PROPELLANT ROCKET TECHNOLOGY

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hours after they have been manufactured or they can be stored indefi­ nitely, always ready when the decision is made to fire.

XI. Conclusion

This summary of solid fuel rocket technology has emphasized the simplicity, reliability, ruggedness, and instant readiness of this type of motor. These are the most important features which recommend solids for greater use in more difficult missions. Other factors are the high den­ sity and long storage life of the propellant and the high mass ratios and thrust-to-weight ratios achievable. Improvements in fuels and greater savings in inert parts' weight promise to increase motor performance and efficiency. The capability provided the designer of solid fuel rockets of choosing a wide range of thrust-to-weight ratios without seriously com­ promising mass ratios provide solid fuel rockets with a patently impor­ tant application in first stages for space applications. Simplicity and ease of ignition make them attractive for upper stage applications.

Bibliography Warren, F. A. (1958). "Rocket Propellants." Reinhold, New York. Sutton, G. P. (1957). "Rocket Propulsion Elements," 2nd ed. Wiley, New York. Zucrow, M. J. (1948). "Principles of Jet Propulsion and Gas Turbines. Wiley, New York. Humphries, J. (1956). "Rockets and Guided Missiles." Macmillan, New York. Huggett, C. (1956). Combustion of solid propellants. In "Combustion Processes" (B. Lewis, R. N . Pease, and H. S. Taylor, eds.), Section M. Princeton Univ. Press, Princeton, New Jersey. Winpress, R. N . (1950). "Internal Ballistics of Solid Fuel Rockets." McGraw-Hill, New York. Summerfield, M. (1958). Burning mechanism of ammonium perchlorate propellants. Part II, Theory of burning of composite solid propellant. Presented at 13th Ann. Meeting, Am. Rocket S o c , New York, November 17-21, 1958. Crawford, B. L., Huggett, C., and McBrady, J. J. (1950). The mechanism of the burning of double base propellants. J. Phys. Chem. 54, 854. Epstein, L. I. (1956). The design of cylindrical propellant grains. Jet Propulsion 26, No. 9, 757-759. Price, E. W. (1954). Charge geometry and ballistic parameters for solid propellant rocket motors. Jet Propulsion 24, N o . 1 (January-February). Newman, R. S. (1955). Solid propellant rocket design. Aero Dig. 71, July, 40-52. Geckler, R. D . (1956). Thermal stresses in solid propellant grains. Jet Propulsion 26, No. 2, February. Arendale, W. F. (1956). Fuel-binder requirements for composite propellants. Ind. Eng. Chem. 48, 725.

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Green, L. Jr. (1954). Erosive burning of some composite solid propellants. Jet Pro­ pulsion 24, No. 1 (January-February), 9. Maeck, A. (1959). Transition from deflagration to detonation in cast explosives. J. Chem. Phys. 31, 162. Dougherty, C. F. (1957). Processing rubber base composite rubber propellant. Chem. Eng. Progr. 53, 489-492. McDermott, J. M. (1958). The role of high polymers in composite solid rocket fuels. Rubber Age 83, 807. Zaehringer, A. J. (1955). Processing composite rocket propellants. Chem. Eng. Progr. 51 (July), 302. Sutherland, G. S. (1958). Solid and liquid rockets—a comparison. Presented at S.A.E. National Aeronautic Meeting, New York, April 8-11, 1958. Hix, W. T., and Bolster, E. W. (1961). Logistic and launch requirements for large solid boosters. Aerojet General Corporation. Presented at Am. Rocket Soc. Space Flight Report to the Nation, New York, October 9-15, 1961, preprint no. 2047-61.

Environmental Control of Manned Space Vehicles ROBERT E . SMITH Department of Physiology, School of Medicine University of California Medical Center Los Angeles, California I. Introduction II. Ecologic Components A. Hypothermic Environments B. Atmospheric Composition C. Atmosphere and Life Systems D. Inert Gases E. Carbon Dioxide F. Water G. Toxic Products III. Bioenergetics of the Steady State IV. Environmental Control: Theory and Applications A. Elementary Theory B. Carbon Dioxide Concentration C. Oxygen Requirements D. Atmospheric Regenerators E. Carbon Dioxide Reduction Systems F. Waste Disposal G. Systems Design—Comparative Aspects References

87 88 88 90 91 93 93 96 97 99 108 109 116 119 120 124 127 130 138

I. Introduction

Essential to the planning for man in space is the recognition that man is of terrestrial origin, and that in the general ecology of man lie impor­ tant sequelae of his Earth-born extraction and gregarious habits. Thus, he reflects in structure and behavior a basic relationship to the polarizing forces of gravity and the periodicities and intensities of sidereal, solar, and lunar events. His concepts of time and space, of life and human des­ tiny arise as the distillate of this diverse continuum of interactions through eons long forgotten in mind yet remembered in his being. And so it may be that a major aspect of man's leap into space may reduce 87

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ROBERT E. SMITH

to the problem of a design of his microcosm so as to include the simula­ tion of sensory inputs essential to his physical orientation and retention of an identification with his earthly self. Apart from energy balance, the prime factors for this may well be simply those of human companionship and, as a likely speculation, an imposed orientational vector simulating that of gravity. II. Ecologic Components

As ecology is the study of organisms in relation to their environment, the ecologic components become intrinsically related within very complex matrices of interlocking life systems. These in turn find their common denominators in terms of mutual requirements for energy and its con­ version into work processes and heat. The basic ecologic components re­ quired of manned space vehicles mainly differ in quantity from those of a well-stocked bachelor apartment or of an igloo. However, there are dis­ crete differences in that space vehicles must present within a balanced unitary system all components necessary for survival. Thus, besides at­ mosphere, food, and varying degrees of reduction and recycling of the products of metabolism, the space vehicle and its biological payload must achieve a balanced energy state compatible with life support in respect to heat storage and quantum exchange. For the balanced energy state to prevail within the manned space vehicle, energy conversions must be provided which are capable not only of meeting power requirements but also of protecting life systems from the irreversible damage of cosmic radiations, be they high energy or high flux density, or both. For such protection it seems that the manned space­ ships of the future will have to carry their own magnetic force fields ca­ pable either of deflecting or of usefully encompassing these bombard­ ments of cosmic energy much as the Earth has done. Also, it would be a fair question to ask geophysicists and cosmologists whether life, as Earth forms know it, could exist on a planet similarly situated to our own but without the corresponding magnetic fields. A. Hypothermic Environments

Because of the linear accumulation of the costs of maintenance energy with time, life support on long missions appears to require either a com­ pletely regenerating cycle or some means of reducing the average energy requirements of the passengers. Even under the assumption of adequate regeneration of food and atmosphere, the long space mission may well call for extended preservation of the life systems on board. The latter suggests

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

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that a minimizing of energy costs and hence a maximizing of survival time could be achieved through such devices as induced cold narcosis, hiberna­ tion, or even preservation in the supercooled vitreous state. Perhaps the ultimate in human null heat states will one day be achieved by dispatching into space the wholly vitrified man to be resus­ citated for duty on arrival at some distant outpost in the deeps of space. Notable from recent work of Audrey U. Smith [1] and associates in Great Britain, the viable reversibility of such supercooled and truly vitreous states has been demonstrated with small mammals. However speculative, such preservation would clearly change, if not reduce, the problems of environmental control and life support in manned space vehicles to something like that now applying to space hardware. A compromise of real interest is, of course, the possibility of inducing the hibernating state in man. This would in practice reduce energy costs of life support in space to approximately 1 per cent of normal metabolic requirements, although still requiring a cabin temperature of some 4 or 5 deg or more above the freezing point of water and a minimal reserve of respiratory gases and atmospheric regenerating equipment. Whether the human hibernator can be developed is unknown; however, suggestions that this may be possible are derived from both anatomical and physio­ logical grounds. If deep hibernation is impossible, there may remain the possibility of inducing transient states of hypothermic torpor, as observed daily in the quiescent bat [2], the hummingbird [3], and in a number of desert ro­ dents [4]. At euthermic temperatures the normal metabolic costs are minimized in sleep, dropping in man by some 10 per cent below the stand­ ard basal metabolic rate. Notably, and irrespective of the temperature level, however, these states would probably not confer additional pro­ tection from ionizing radiations. In addition to energy balance, ecologic considerations govern the totality of inputs, outputs, and interactions of the life support compo­ nents. A number of these will be considered under such categories as at­ mospheric composition, body metabolism and temperature regulation, water and food balance. Within this framework, but outside the scope of this chapter, are a variety of other factors of the space environment such as null gravity, magnetic flux, minimal sensory inputs (including those normally present in human associations), and the terrestrial diurnal rhythms. Related to these are a number of corollaries derived from the physical properties of matter. Among them is the obvious fact that movement of fluids, whether gaseous or liquid, becomes in space almost entirely de­ pendent upon forced convection processes unaided by the usual density

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gradient factors. In lieu of this, the mixing of gases becomes solely de­ pendent upon vapor pressure and pure kinetic diffusion processes. Similarly, uncertainties exist about the disposition of the relatively immiscible materials collected within the microcosm of the living cell. Here we may not rely solely upon the behavior of such time-honored creations as the egg of the sea urchin, but also must examine the cellular systems of mammals wherein every cell is bathed in the nutrient stream of a cardiac-driven convective transport system that is itself sensible to the ebb and flow of demands upon total body work (cf. [51]). Again, we need to know whether the flatworm's (Planaria) interesting nucleic acid coding system can retain its configuration and mnemonic devices against the composite environment of null gravity, minimal magnetic flux and the cosmic radiation flux of outer space. I t is for these and related basic reasons that the National Aeronautics and Space Administration (NASA) is now planning to place a number of purely biological payloads into re­ coverable space probes. These same questions are being posed on a somewhat more complex level of organization in highly technical experiments on small mammals and primates that are sent for protracted trials in space. These studies reflect the serious concern that has arisen over the possibility that the ecologic circumstances of extended exposure to the space environment will fail to furnish sufficient sensory inputs. These sensory inputs must maintain in the brain stem the minimal electrical activity compatible with functional survival of the higher brain centers and the association areas of cerebral cortex [6, 7 ] . All of these, and many more such ques­ tions will have to be answered before man ventures into deep space. Per­ haps most certain is the probability that when man achieves his optimal environment in space we will come to know far better not only his true nature but his true place in nature. B. Atmospheric Composition

Specifications on atmospheric composition of space capsules have varied greatly between the extremes of pure oxygen at a po 2 °f around 258 mm Hg (5 psi) to one of pure country air at standard atmospheric pressure. Certainly, for short missions such as that of the Mercury manned satellite program, the pure oxygen system at 5 psi has proved reasonably successful, and there is now evidence [8] that men can toler­ ate pure O2 at a po2 of 250 mm Hg for a period in excess of 16 days. Among the arguments favoring a reduced total pressure are those indi­ cating lower outboard leakage rates from the vehicle, whether from in­ trinsic or extrinsic causes, such as meteorite perforation of the hull [9].

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It is also true that rapid decompression from a pressure of one at­ mosphere of air to the equivalent of 4-5 psi would invite the hazards of decompression sickness ("bends") due to excesses in the partial pressure of N 2 dissolved in the body tissues. As a reasonable compromise it has been suggested that air at a total pressure of about 500 mm Hg (11,000 ft elevation) would prove most nearly optimal [10] for longer periods of flights into space. Compared with a full atmosphere, this level of total pressure (with p02 at 150 mm Hg) would extend the emergency time available on hull perforation, would reduce somewhat the intrinsic leak rate, and, in the event of reduction to a standby pressure of 5 psi on pure oxygen, would protect the astronaut indefinitely from the "bends." For these and additional reasons relating to physiological requirements and thermal balance within the comfort zone, it has been proposed [10] that the atmospheric composition be held as follows: total pressure (in mm Hg) = 500; 0 2 = 150; C 0 2 = 2-4; H 2 0 = 14; N 2 = 334; air temperature = 26°C. C. Atmosphere and Life Systems

The argument favoring a basic composition of atmosphere similar to that close to Earth's surface is, in many respects, supported more by our ignorance than by well-established information. However, there are sev­ eral lines of evidence which appear to converge on the affirmative side. Among these is the evolutionary history of the life system itself. If we assume, as do most biological scientists today, that life arose de novo on this planet at a time when Earth was an essentially anaerobic (i.e., oxy­ gen free) void of rock and water, it appears safe to postulate that the present tropospheric atmosphere attained its present composition by life systems which developed through eons of time. Aiding in this, however, was also the slow but steady yield of oxygen derived from the photo­ chemical dissociation of water vapor entering the upper mesosphere where the resulting hydrogen escapes into outer space (cf. Bates and Nicolet [11]). As Kaplan has noted [12], this occurs at rates sufficient to gen­ erate in a billion years one-tenth of our present atmospheric oxygen. By retrospective reasoning based on present knowledge of the chem­ istry of the living cells of animals and plants, it may be deduced that what is now the cell nucleus was derived under an essentially oxygenfree atmospheric environment. In short, this means that the early forms of life received their energy from quantum utilization processes not re­ quiring oxygen as the terminal electron acceptor. Parenthetically, this stage of life may be regarded as the pioneer breakthrough to "Advances in Space Science," of which perhaps a neonatal phase may again be ap-

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pearing with the human creation of the solar battery. In any event, it is quite clear that the life system was also to become a principal agent in supplying the chemical potential necessary for the derivation of an oxy­ gen atmosphere from the original terrestrial reservoir of bound oxides, including water. Doubtless the photosynthetic process, as it developed, also greatly accelerated this oxygen liberation at a rate far exceeding the initial rate. On this assumption the curve of atmospheric oxygen concen­ tration with time may have presented a sigmoid shape with perhaps an inflection point in very early Pre-Cambrian times. The theory follows that as atmospheric oxygen levels became gradu­ ally elevated, the life systems which were to persist became faced with a variety of new problems. Among these the most serious was, in a sense, the protection of the replicating machinery from direct oxidative degra­ dation. This was evidently solved through the provision of an electron transfer system by which the heretofore anaerobic life system could suc­ cessfully cope with an aerobic environment. In effect, this is what now surrounds the cell nucleus in all of the classical forms of life on our planet. However, as Herbert Stern pointed out some years ago [13], the nucleus still remains essentially an anaerobic system, surrounded by its cytoplasmic envelope. In the cytoplasm are contained a host of supportive functional activities, principal among which is the biological machinery for harnessing to cellular requirements the energy derived through oxi­ dation of hydrogen by atmospheric oxygen. While the biologists may quibble over the means by which the cell has evolved, there is general agreement on the empirical fact that the cell nu­ cleus appears quite sensitive to oxygen and may be readily damaged by high values of po 2- This is probably because it has intrinsically no satis­ factory electron donor systems by which it can couple relatively non­ essential hydrogens to oxygen. Paramount here is the fact that oxygen has the effect of depriving sulfur of the hydrogens necessary to maintain certain S-H (sulfhydril) bonds critical to normal function of the nuclear proteins. A similar situation threatens in the case of the sudden appear­ ance of large amounts of oxidants within the vicinity of the nucleus. Thus, high levels of penetrating ionizing radiation result in nuclear dam­ age of this type, in addition to the target-hit effects upon the chromo­ somal structures. In addition, the oxygen-utilizing particles within the cytoplasm (i.e., mitochondria) are also heavily, but transiently, damaged by ionizing radiation, though it is not yet clear whether these are pri­ mary results of target-hits or are secondary to the local oxidant concen­ trations induced by the ionization of the oxygenated cellular water [14]. The net effects of high oxygen levels, therefore, are evidently to saturate

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

93

the cytoplasmic capacity for terminal electron transport, and in the vi­ cinity of the nucleus to oxidize S-H and possible other bonds essential to the functional integrity of the system. D. Inert Gases

As atmospheric nitrogen is again a heritage (though scarcely a prod­ uct) of the life environment, one may reasonably ask whether this gas is really biologically inert and serves only as a diluent against the threat of oxygen toxicity. To this question there is not a clear answer. In experi­ mental situations it is often difficult, if not impossible, to distinguish be­ tween effects arbitrarily assigned to oxygen toxicity or to lack of nitrogen (or to both). Allen has recently reported [15] that embryonic tissue in vitro fails to grow normally in a mixture of gas which at one atmosphere does not contain more than 10 per cent nitrogen. Earlier Cook [16] showed that for a given po 2 the total metabolic activity of various ani­ mals was lowered whenever the total pressure was reduced below one atmosphere. Could it be in the case of the human premature infant that the retrolental fibroplasias ascribed to the use of a pure oxygen atmos­ phere are in some degree due actually to a lack of nitrogen [15]? Similar questions come up in respect to the pulmonary irritation ob­ served from extended breathing of pure oxygen at sea level. Furthermore, in diving experiments, nitrogen as well as all of the noble gases (except­ ing possibly helium) become toxic as their effective concentration is raised. Of these, xenon at 80 per cent in oxygen at sea level is a good anesthetic, while nitrogen in excess of 4 atm or more induces symptoms of narcosis [17]. Since it is evident under these circumstances that these "inert" gases are really not inert biologically, one may not exclude the possibility that our present life systems are perhaps optimally suited by a p N 2 of approximately present atmospheric levels. E. Carbon Dioxide

Carbon dioxide, while conceptually much maligned as a toxic meta­ bolic end product, is again not as much the enemy as it is the friend of men and animals. A prime benefactor through the photosynthetic cycle of plants, this agent becomes the universal vector of carbon transport as water does in respect to hydrogen. Beyond this, however, C 0 2 , as the highly water soluble acid anhydride of carbonic acid, determines pri­ marily the acidity of virgin waters and also provides the major reservoir of labile hydrogen ion buffer material in the body fluids, particularly of

94

ROBERT E. SMITH

terrestrial organisms. This becomes more evident from the combining re­ action and resulting ionization in water C 0 2 + H 20

= H2CO3

(1)

H 2C 0 3 - H+ + HCO3-

+

wherein the acidity ( H ) of the water is clearly increased as a function of atmospheric partial pressure of C 0 2 (Pco2) (2)

The significance of this relationship to the blood chemistry of man in a closed system is more apparent when we recall that with pH = —log(H+) and pK = —log k, the Henderson-Hasselbach equation then gives us (3)

Since normally in the blood the ratio of bicarbonate concentration sl ( H C 0 3 ~ ) to that of carbonic acid (pco2) °f the order of 20:1, the buffering power of bicarbonate in blood allows for a considerable increase in C 0 2 concentration without greatly changing the pH of the blood. One rather common problem emerges, however, whenever the subject at rest overbreathes enough air to wash out the body C 0 2 at rates exceeding those of production. In this case the pH tends rather readily to shift to­ ward the alkaline side, and a so-called hypocapnic (low C 0 2) condition is induced with accompanying alkalosis marked by peripheral spasticity and superficial cyanosis, followed later by varying degrees of uncon­ sciousness and an abnormal (Cheyne-Stokes) breathing pattern. Thus, not only is tolerance for high environmental C 0 2 effectively greater than that for lowered blood C 0 2 , but also at the higher ranges of physiological e values of pCo2 ( -g-> 40-50 mm Hg) the oxygen-carrying capacity of the blood hemoglobin becomes significantly increased. In a sense this consti­ tutes a further category by which C 0 2 serves in the cause of animal sur­ vival. Finally, as suggested earlier, the role of C 0 2 as a vector of carbon transport enters prominently into the normal metabolism not only of plants, but of animals as well. In both it enters into numerous exchanges involving synthetic as well as degradative reactions within the cell. From these sweeping allusions one may at least gather that CO2 is as necessary as water as a component of the life system and that it is an adverse lim­ iting factor only when, like water or oxygen, its concentration rises or falls beyond the compensatory capacities of the organism. Maintenance of carbon dioxide control becomes a critical factor in the space capsule in both the biological and engineering design aspects. The

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

95

partial pressure of C 0 2 must be kept at a level that permits satisfactory unloading of the venous supply of C 0 2 presented to the lung alveoli. At any given metabolic activity the unloading rate is normally kept con­ stant by the combined effects of lung ventilation rate and the pressure a n differential between the pCo2 of venous blood ( = alveolar pco2) d that of the ambient atmosphere breathed. For chronic exposures the pGo2 of the systemic arterial blood should be held within its normal range, as represented by alveolar levels of from 30 to 48 mm Hg (mean = 36.7 with S.D. = 2.7 for 186 observations on air at PB = 733 mm Hg) [18]. In steady states the normal range is ordinarily maintained by reflex com­ pensatory adjustments, so that within limits of ambient pco2 from 0 to 32 mm Hg, the lung ventilation increases from normal levels of around 8 to 10 liters/min to about 20 liters/min. Beyond the value of pCo2 = 32 mm Hg, the unloading of C 0 2 quickly becomes possible only by an increase of blood pCo2 whereby a differential of from 7 mm to 2 mm Hg between blood and alveolar pCo2 is achieved up to a C 0 2 level of the order of 70 mm Hg. Effective elimi­ nation of C 0 2 at such levels may be achieved only by an enormous in­ crease in lung ventilation rate. Above the critical level of pco 2 = 32 mm Hg, the lung ventilation rate rises approximately as a straight line func­ tion of the ambient C 0 2 concentration to reach the order of 80 liters/min at pco2 = 75 mm Hg. The practical conclusion to be drawn from these relationships is that Pco2, ambient, should not be allowed to exceed an upper limit of about 30 mm Hg, at least under conditions of normal po2 levels. It may be noted in passing that a pCo2 slightly higher than this can be satisfac­ torily tolerated under transient periods of hypoxia, where a moderately raised pCo2 appears actually to improve performance. In human experi­ ments at ambient po2 of about 75 mm Hg, an optimum combination of ventilation rate and arterial po2 was achieved by raising the ambient Pco2 to the order of 23 mm Hg [19]. Apart from vasodilatory effects, a large measure of this advantage is evidently conferred by the "shift to the right" effected upon the blood oxyhemoglobin dissociation curve by raised C 0 2 . This results in a greater unloading of oxygen from the blood as it perfuses the tissue capillary beds. None of these devices should be employed, however, except possibly for emergency purposes, since any increase in blood pCo2 is achieved only at the cost of increased blood acidity. In chronic exposures the lat­ ter would carry the concurrent threat of serious displacements in the electrolyte balance of the body. Deleterious effects of long-term exposures to high C 0 2 levels and various oxygen tensions have been described by Schaefer [20] in a report on submarine atmospheres. Effects included

96

ROBERT E. SMITH

evidence of increased blood supply to the skin, impaired cardiovascular response to exercise, a fall in mean arterial pressure, and a reduction of oxygen consumption. In addition, the men showed an impaired attentiveness as measured by the Burdon test. F. Water

Water and water vapor, primary among biological essentials, present many problems having common roots in both design and physiological function in closed spaces. A later section will demonstrate how the control of water vapor is closely linked to that of body heat and fluid balance, and the relation of these to the regulation of ambient temperatures through the evaporative and convective transfer of heat. Furthermore, in achieving recycling sufficient to meet the demands of daily fluid in­ take, the space capsule design must also embrace the important facts, namely, that water is itself a metabolic end product and that the stored food supply is converted about equally to both C 0 2 and H 2 0 . Thus, water stores will normally accrue in direct proportion to both metabolic turnover and time and, like C 0 2 , with no appreciable change in the total mass of the system. Total daily water exchange for human adults under conditions of average temperatures and humidity is approximately as follows: Water Intake

(ml)

Water Output

(ml)

Dietary solids Oxidative wate Fluid ingestion

1200 300 1000

Feces Urine Respiratory Skin

150 1500 350 500

2500

2500

Since the daily water derived from oxidation of food ("metabolic water") averages around 14 gm/100 cal of dietary intake, the net increment to the water storage pool would amount to some 300 to 350 gm/day. The sum of respiratory and skin losses adds up to nearly the total fluid ingestion requirement for man. This suggests that with an efficient recovery of these two sources alone, the additional water required for drinking purposes would be of the order of 300 ml/day. Certainly for short-term operations this order of need could be better stowed than ob­ tained by reclamation from urinary sources. However, for the longer term need, Sendroy and Collison [21] have obtained from urine a wholly

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

97

potable water in 99 per cent yield through lyophylization and a single passage of the condensate through activated carbon. Of obvious concern are the subordinate effects of water vapor upon many features of the physical system, e.g., electronic equipment, fogging of transparent surfaces, condensation and icing of areas of low tempera­ ture and high heat conductance, occlusion of particulate regenerator com­ ponents, etc. One physiological aspect of currently indeterminate nature arises from the finding in recent years that postural changes profoundly in­ fluence urinary output. This is done through a reflex exaggeration of gravity-induced thoracic blood volume changes. These inhibit urinary output during standing and accelerate it during recumbency. However, in a null-gravity space system our subject remains constantly in a state of relative inactivity and recline. The question may be raised as to whether the normal antidiuretic mechanisms will function effectively to maintain steady state water balance under prolonged exposure to these conditions. G. Toxic Products

The approach to atmospheric toxicology of space vehicles will prob­ ably remain more or less empirical despite the desires of everyone con­ cerned. Insofar as possible, each new component of the closed compart­ ment must be checked for volatility. This must be checked as a function of several sets of physical factors with ranges of intensity and duration that vary with the vehicular design, the life support system used, and the nature of the mission. The same factors of uncertainty may be applied currently as were outlined by Specht [22] a decade ago. Basic principles have been long established with industrial hygienists. They include exclusion from the compartment of all volatile liquids, particularly those such as hydrocar­ bons and organic solvents of high vapor pressure and known toxicity. Equally critical could be the piping and valving (or both) of such ma­ terials within the life support spaces. The hazards of high temperature effects are serious in terms of either inboard electronic gear or exterior atmospheric heat effects. Because of the possible cumulative effects of atmospheric contaminants at very low concentrations, it is highly important to maximize the specific surface of the adsorbents and to give careful attention to possible occlu­ sion or "poisoning" of the catalyst or of the adsorbent columns. Sweepout rates are basically linear in their concentration dependence; there­ fore, the effective clearance fraction is always a function of the amount

98

ROBERT E. SMITH

remaining to be cleared, i.e., the amount removed per cycle is a fixed per­ centage of the amount remaining. However, depending on the adsorbent properties this may be effective only down to some concentration at which the amount desorbed may equal the amount adsorbed. In any anticipated circumstance full assurance of effective clearance must be repeatedly tested by empirical trial. Another hazard in the space vehicle is the effect of ionizing radiation upon the enclosed atmosphere. As Schaefer has reported [20] in his stud­ ies on submarine compartments lined with radium-painted dials, the con­ centration of positively charged air particles rises in proportion to the radiation intensity. These particles may have an irritating effect upon exposed mucous areas; further research on their properties is indicated. Toxic products may be derived from the biological as well as the physical components of the system. In normal man the nutrients of food are absorbed mainly from the small intestine where, besides water reabsorption, a great deal of degradation of bacterial origin is carried on. The end products of this activity (cf. [23]) include such gases as C 0 2 , CH 4, N 2, H 2S, and organic acids such as acetic, butyric and lactic acids. Bacterial decomposition of lecithin is of particular interest in giving rise to choline and to related amines of a toxic nature including neurine and muscarine. Likewise, various amino acids are decarboxylated to give toxic amines (ptomaines), including also histamine, putrescine, tyramine, cadaverine, and agmatine. The odor of feces is derived from indole and skatole resulting from degradation of the amino acid tryptophan; H 2S and methane as well as various mercaptans are normally formed from cystine. Skin secretions generate small daily amounts of organic acids [24] including capronic, caprylic, valerianic, and butyric acids be­ sides acetic, lactic, and formic acids. Ammonia, also formed in the large bowel, is reabsorbed into the blood to appear eventually in nonvolatile form in the urine. However, under certain conditions of ketogenic acidosis, N H 3 may appear in greater amounts in the urine; in standing urine, urea may give rise to N H 3 through degradation. During ketosis, acetone may appear in both the expired air as well as the urine. The asparagus odor of urine is probably due to a methyl mercaptan. Further study is needed on the effects of chronic exposure of men to low levels of these volatile biological end products. Perhaps fortunately in this instance, the threshold of smell is fantastically low and so provides man with a built-in monitor for many of these potentially highly-toxic agents. However, in respect to toxic agents of an odorless nature, several categories of volatile contaminants may require evaluation for toxicity and tolerance levels. Moreover, it is now known that many organic compounds, including

99

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

those appearing in flatus (cf. [25]), cause serious contamination of ad­ sorbents such as those of the recently developed "molecular sieves" (cf. Willard [26]). Special trapping of such agents will be requisite to the recycling methods employing these adsorbents for regulation of C 0 2 and water vapor. Notably in the Mercury capsule a reportedly [27] success­ ful elimination of odoriferous agents has been achieved through an acti­ vated charcoal absorber inserted in the line ahead of the LiOH C 0 2 absorbing chambers. In the dog-carrying Soviet Sputnik 2, odors of urine and feces were stated to be well controlled by inclusion of a specially dried absorbent moss and activated charcoal in the receiving receptacle [28]. However, the strongly oxidizing K 0 2 (?) of their regenerator may have been contributory as well (cf. [29]). III. Bioenergetics of the Steady State

An operational definition of "steady state" may be expressed simply as one in which the energetic inputs and outputs of a given system are equal. For man in space we can arbitrarily require that the body metab­ olism be equal to the sum of heat and work outputs plus the caloric equivalents of any storage function. The latter may include changes in weight or composition of the body as well as actual gains or losses in the heat content of the body. In such a system mass remains constant but energetically its enthalpy degrades irreversibly according to the Second Law unless extrinsically supplied with inputs equivalent to the entropic losses. As a basis for reference, an operational equation may be written as Qm =

Qheat +

#work,

(4) 2

where qm (or M) is the metabolic heat production in kcal/meter /hr; q W t will include all heat terms whether of heat storage or loss, and tfwork is the caloric equivalent of external work. In the ordinary "basal metabolic" rate of human subjects, the work term is arbitrarily reduced to zero by keeping the patient under standardized conditions of rest, while the heat terms are further minimized by having the patient in a 12-hr fasting state and appropriately maintained at a thermo-neutral temperature and humidity. Under these conditions the measured value of qm represents the oxygen cost of body maintenance under "basal" con­ ditions. Reasonably accurate formulae are available upon which one may predict metabolic costs from a knowledge of body size and composition. With normal subjects, major departures from standard prediction formu­ lae have been experienced mainly in individuals whose fat content ex­ ceeds the normal proportions, i.e., 10 to 16 per cent, respectively, for men

100

ROBERT E. SMITH

in the age range 25 to 45 years. For women the values are 15 to 23 per cent [30]. For this range the basal metabolic heat production is given for 2 a variety of body types as: 40 to 42 cal/m /hr for young men, and 36 to 2 37 cal/m /hr for women. Body surface in square meters may be computed from height-weight data by such formulae as that of Dubois, i.e. 2

6

5

(surface area in meters ) = TF&g X ff&2> X 0.007184 2

(5)

giving respective values between 1.6 and 1.8 meters for the average women and men noted above. The usual conversion factor taken for the caloric equivalent of oxy­ gen, i.e., 4.83 Cal = 1 liter 0 2 , assumes that the respiratory exchange quotient, i.e., the ratio, C 0 2 production/0 2 consumption, to be 0.83; about 250 cc 0 2 per minute (STP) are required for a man that weighs 150 lb. Although data suitable for general specifications on respiratory re­ quirements are obtainable from a number of metabolic reference tables [31], it is well to repeat that any long-term evaluation of the needs of a given individual should be obtained from direct measurements on the subject under various conditions, both in basal state and under work loads of the expected range required of the operation. The latter are particularly important in evaluating maximal requirements for gas ex­ change in a closed space, since under conditions of heavy work loads the oxygen requirements and heat production may rise to ten or more times the resting levels. Concurrent increases in per minute lung ventilation volumes may develop from the normal 8 to 12 liters to 50 liters/min, with peak tidal excursion velocities of the order of 125 liters/min. Physiologically, life in the null-gravity state faces its own peculiar problems in the maintenance of normal functional capabilities. Thus, while we have long been aware of the debilitating effects of prolonged bed rest, it has been emphasized in recent years that such inactivity re­ duces the mass of the heart as well as skeletal muscle. In the space en­ vironment as elsewhere, a new adjustment may eventually be reached consistent with the reduced work load imposed upon the body; however, this residual level of work capacity may be incommensurate with the output required to sustain normal posture and locomotion against ter­ restrial gravity or for a work load. In space, man will not be obliged to perform much work in the clas­ sical sense, since his energies will be devoted largely to the exertion of essentially conservative forces. Considering this, and in the absence of the usual sensory inputs leading to muscle tonus and positional orien­ tation, there may be little use for many of the postural righting reflexes, or for the usual tonus of the antigravity musculature. Hence, physical

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

101

deterioration under these conditions may pose a threat to the capabilities of the future astronaut in space sufficient to demand counteraction through some carefully designed regimen of exercises on spring-loaded machines. The capacity of the adult human body for maintenance of the "steady state," i.e., homeostasis, is no more elegantly exhibited than in the regu­ lation of water balance and body temperature. Water is the prime solvent and medium of solute transport in the system as well as the heat sink and heat exchange fluid. Thus water economy is an essential common denomi­ nator to both the energy balance and the atmospheric control of the space cabin. As shown in Fig. 1, the body heat dissipation becomes com­ pletely dependent upon evaporative cooling as the ambient air tempera­ ture rises a degree or so above the mean skin temperature (33° to 34°C). A second point to note is that the comfort region of air temperature lies between about 18° and 28°C and, largely by reduction in clothing, may be shifted toward the upward range in warm summer weather. Chronic exposures at either extreme may be attended by a varying degree of phys­ iological acclimation involving changes in metabolic heat production and in water and salt balance. The various environmental conditions of air movement, humidity, temperature, and barometric pressure governing the range of human tol­ erance and performance have been evaluated largely through direct em­ pirical studies. These have entailed an enormous amount of arduous and often heroic experimentation (cf. [32-39]). For reduction of this complex response pattern to theoretical applica­ tion in human engineering practice, we are indebted to the late Professor Craig L. Taylor and his associates in the School of Engineering at the University of California at Los Angeles. Because of his work, the prob­ lems of atmospheric control of water vapor and ambient temperature within a closed compartment can be approached in a far more rigorous fashion than has heretofore been feasible. In this chapter, however, it is proposed only to examine some of the more important physical parame­ ters in relation to physiological requirements. Certain general formula­ tions and examples will be offered, but more for instructive purposes than for explicit description of actual design specifications. Adopting the notation of the above group (i.e., Taylor et al. [cf. 36]), we may expand Eq. (4) to give an equation of heat balance qs = qm — qr ~ qc — Qei - qe*

(6)

in which q8, the rate of heat storage, is given by the difference between metabolic production, qm, and the heat exchanges, respectively, by radia­ tion, qrj convection, qc, and evaporation through insensible, qei, and sen-

102

ROBERT E. SMITH

sible, qe8, routes of exchange. The practical problem in thermal regulation is to set limits to the ranges of qs which will be acceptable under speci­ fied operational conditions. For performance at physiological steady 75°F

82.5°F

100

91.5°F

94°F

PER CENT HEAT LOSS

['.EVAPORATION

23 30 AMBIENT AIR TEMPERATURE

35°C

FIG. 1. Distribution of body heat losses through vaporization, convection, and radiation as functions of ambient dry bulb air temperature (after Brody [ 3 1 ] , from basic data of Hardy and Dubois [ 3 3 ] ) . The vertical lines indicate, respectively, the approximate temperature ranges of the winter and warm summer comfort zones, and the mean skin temperatures of resting and exercise states (see Yaglou in [ 3 8 ] ) . The mean skin temperature of the threshold for visible sweating is denoted by the line, 8, and horizontally, by the corresponding air temperature range.

state, qs = 0; ideally this should be achieved at minimal metabolic cost and in an atmospheric environment within the thermal comfort zone. However, this may well prove impossible during transient periods of high

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

103

thermal gradients such as those of reentry or on certain other trajectories within the ionosphere. In these cases the limit will be set by the time-rate integral of heat storage compatible with retention of operational capa­ bilities essential to completion of the mission [36]. This principle has proved valuable in application to supersonic flight within the strato­ sphere. For missions in space the assumption either of thermal overloading or of net body cooling should be held strictly secondary to a satisfactory solution holding qs = 0. In the latter case we then have qm = qr +

qc +

qi +

qe8

(7)

from which qe8 can also be eliminated, at least during inactivity, by spec­ ifying an ambient temperature below the sweating threshold, e.g., ta = 26°C and water vapor pressure, pa, 12 mm Hg. The insensible evapora­ tion rate, qei} is a direct function of total barometric pressure as well as atmospheric water vapor pressure [32]. The empirical regressions ob­ tained by Hale, Westland, and Taylor [34] on resting man give for total 2 evaporation loss (gm/meter /hr) We = 29.57 - 0.012 pB - 0.334 pa, S.E. = ±1.10

(8)

of which insensible loss via the skin is W8 = 21.15 - 0.012 pB - 0.117 p«, S.E. = ±1.14

(9)

These regressions hold for nude resting men breathing oxygen over the experimental limits of PB, 760 to 253 mm Hg and paj 6 to 26 mm Hg; the ambient air temperature is 27.7°C and the mass velocity is 558 k g / 2 meter /hr. A computation of qei from Eq. (8) for PB = 600 mm Hg and pa = 12 2 mm Hg would predict an insensible total water loss of 18.4 gm/meter /hr, 2 which for a skin temperature of 34°C would amount to 11.2 kcal/meter / hr or some 23 per cent of M, the resting metabolic heat production rate. For illustrative purposes the residual of Eq. (7) may be evaluated as qr + qc = qm — (qei + qes)

(10)

= M - 0.23 M = 0.77 M, if qes = 0 Solutions of qr and qc in terms of temperature differences and conduct­ ances (cf. [34, 36]) give qr = 3.59 (ts - tw)

(11)

and qc = 0.0735 G

05

(T/To)°'» (ts - ta)

(12)

where subscripts refer to s, skin; w, wall; a, ambient temperature in °C;

104

ROBERT E. SMITH

and T is either skin film temperature (°K) or an operative temperature averaging the surface and ambient temperatures. Combining these in Eq. (10), we have 0 5

05

0.77M = 3.59(*. - tw) + 0.0735 G - (T/T0) '

(ts - ta)

(13)

2

G (kg/m /hr)

from which various relationships may be derived as needed. At this point the relation of the respective parameters in Eq. (13) to

3

0.2'10

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

AIR VELOCITY V*km/hr

FIG. 2. Sea-level mass velocity versus air velocity required of various densities of moist air for indicated temperatures and atmospheric pressures.

the heat economy and water balance of man in the space compartment may be established. Of key importance to environmental control is the definition of G, the sea level mass velocity term, as a function of the den­ sity, D, of moist air and the velocity, 7*, of air movement, i.e. G = [Do(T0/Ta)(PB

- 0.3783p«)/P 0]F*

(14)

It is evident that in this expression are factors relating both to thermal regulation and to water vapor, as well as to barometric pressure and, by inference, to respiratory gas exchange (see below). Since the enclosed radical is explicitly the density, D, we have simply

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

105

G = DV*

(15)

For later application in design we will need to specify values of each term, as both D and V* have critical limits and their ratio is everywhere uniquely defined for a given value of D (Fig. 2). Thus the density term expresses the roles of ambient temperature, total pressure, PB, and water vapor pressure, pa, in determining the relative values of G and V* and thereby the thermal regulation through convective conductance. From the definition of convective conductance, i.e., [36] as 5

0 5

hc = (r a/r 0)°' (0.0735 G - )

(16)

it is evident that h0 rapidly approaches a vanishing point as G is per­ 2 mitted to decrease below a value of around 100 kg/meter /hr (Fig. 3).

2.0

1.8

1.6

1.4

1.2

1.0

0.8

0.6

0.4

0.2

16

200

400

800

2

G (kg/m /hr)

FIG. 3. Convective conductances, calculated according to the equation, are given for various temperatures and mass velocities.

106

ROBERT E. SMITH 2

An arbitrary level of G > 400 kg/meter /hr is required. Failure to achieve this will virtually abolish convective conductance and result in throwing the burden of heat exchange very greatly upon radiation con­ ductance. For the latter the coefficient is some 50 to 80 times larger than 140

120

100 80

60 ta(°C)

40

20

0

-20

-40

-60 -40

-30

-20

-10

0

10

20

30

40

FIG. 4. Graphical representation of role of mass convection velocity, G, in deter­ mining extent of thermal regulatory stability. Higher values of G produce greater 2 stability. Note that increments of metabolic heat production, M(kcal/meter /hr), do not alter the slope. Values of ta were calculated from Eq. (16) for arbitrary values of G and tw as shown.

that for hc. Further, the heat transfer now depends upon the difference in temperatures between the skin or garment and the wall (Eq. (11). The resulting effects upon ambient air temperature are easily illustrated by solving Eq. (13) to give ta = [3.59(*. - tw) + Kts - 0.77M][1/AJ

(17)

107

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

Substitution of numerical values into the right hand side of this equation gives a set of solutions for ta as a function of tw for various values of the conductance, hcj as obtained from Eq. (16) or from Fig. 3. Thus, plotting ta versus tw (Fig. 4), we observe that the very steep slopes obtained with small values of G are greatly reduced as G is increased to 400 or more. 2 The effect of increased metabolic output from 49 to 80 kcal/meter /hr is 300

200

y=

96 .5 - 3. 59 t

100

c

3 -loo

w

e

-200

-300

-400

-5001 -40

-20

20

40

60

80

100

120

140

160

t °C

FIG. 5 . The curve of heat increment, 2/(kcal/meter/hr) beyond metabolic heat, necessary to maintain a constant ambient air temperature of 2 6 ° C , as a function of wall temperatures. The equation is derived by substitution of appropriate numerical value's into Eq. ( 1 3 ) .

also shown in this figure and illustrates the fact that heat production within the system will not change the slope of the curve for ta versus tw unless associated with a change in conductance values. The question of the maintenance of steady state temperature and thermal balance within the closed vehicle is now examined. Using Eq. (13) we may set ta constant at 26°C and require that supplementary 2 kcal/meter /hr be introduced as wall temperatures are lowered. This pro­ cedure leads to a straight line (Fig. 5) of slope determined solely by hn

108

robert

e.

smith

the radiation conductance. Here the convection terms contribute only insignificantly to the constant denoting the intercepts, whence no appre­ ciable advantage accrues in this instance through increases in G. This, however, does not mean that convection conductance may be disregarded, since the latter continues to determine the stability of the thermal regu­ latory characteristics as noted earlier. The conclusion from Fig. 5 is that we will have to add or subtract heat from the system at the rate of 3.59 2 kcal/meter /hr per °C change in wall temperature* whenever the latter is, 2 respectively, below or above 27°C. For the standard man of 1.8 meter surface, this would amount to about 6.5 kcal/hr/°C of change in tw and, assuming no change in skin temperature, would be independent of the steady state air temperature selected.

IV. Environmental Control: Theory and Applications

Earlier sections were concerned chiefly with man as a biological sys­ tem possessed of specific requirements for materials and exchanges of energy entailing metabolic activity and performance of external work. It was also indicated that the physiological machinery for this depends directly upon the maintenance of steady states within the internal en­ vironment of the body and that in respect to the external environment the exchanges of food, water and respiratory gases as well as work proc­ esses are also attended by continuing degradation of these materials and activities into heat. The present section provides a basic framework of theory for specifiying the characteristics of an environmental control compatible physiologically with requirements within the manned space vehicle. From what is already known of such requirements, this problem may be regarded as one of optimizing a recirculating gas system in order to provide adequate atmospheric regeneration concurrently with proper heat transfer and thermoregulation of the compartment. In such a model the physiological requirements dictate both atmospheric composition and gas concentrations as well as rates of regeneration or renewal of various components, or both. From a solution to this we may then specify both net compartment volume and ventilating characteristics appropriate to the physiologically required mass velocity. * Since the value, hr = 3.59, assumes certain shape factors for the enclosure and a body radiation area of 64 per cent of body area, the actual heat losses could exceed those predicted here by possibly twofold, depending upon the selection of the emissivities and shape factors in the design.

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

109

A. Elementary Theory

It is assumed that for a working model the biological component is exchanging an tth gas, i.e., carbon dioxide, nitrogen, water vapor or other gaseous components within a closed system of volume of V (liters). Gase­ ous elimination products, mostly C 0 2 and H 2 0 vapor, are being produced continuously at respective rates of Q mols per min. These products are assumed to be removed at a rate proportional to their concentrations d (mol/liter), i.e., at a rate ktd; here kt is an em­ 1 pirical removal constant (dimensions t" ) given arbitrarily as fc* = f(r/V, cii) and in practical terms as ki = rcii/V, with r the pump circula­ tion rate (liters per minute), and a* a dimensionless constant designating the absorbent or scrubbing characteristics for the gas involved. On the assumption, initially, that V, Qiy and ki are constant, the equation of material balance* may be written as dd/dt

= (Qi/V) - hd

(18)

Then integrating (19) whence at t = 0 (20) where Cio is the initial value of C*. Solving for d gives (21) or (22) and also d = (Qi/kiV)(l

kit

- e~ ) + de~

ka

(23)

Equation (23) conveniently expresses the fact that the concentration of the ith gas, Cif is proportional to the rate of production, and inversely proportional both to the removal constant, kiy and the total volume, V. * For the solution of this set of equations, the author is pleased to acknowledge the kind assistance of Dr. George E. Moore of the Department of Physiology, UCLA Medical Center.

110

ROBERT E. SMITH

It also serves to emphasize that C% can be treated as the resultant of two operational components, i.e. (1) a term expressing the initial conception, CiQ, which decays exponentially, and (2) a term asymptotic in time to Qi/kiV. Significantly, this means that as t becomes large, d approaches Qi/kiV independently of the initial concentration, d, Fig. 6. Furthermore, 10

Qi/k.V= 1.0=0.

o C. = 10

3

• e.

o

= 0 . K Q . / k l.V

~~

7

6

4

3

2

1

ki V 1.0 t-At

t

1

2

3

4

5

6

TIME

FIG. 6. Plots of gas concentration versus time, showing that the same asymptotic values are reached in the same time whether initial concentration is greater or less than that of the specified asymptote.

the curve of Ci(t) displays neither maximum nor minimum, providing, of course, that no changes occur in the magnitude of Qi/kiV. For ki con­ stant, the time required to approach this asymptotic value is also inde­ pendent of the time at which the system is started.

E N V I R O N M E N T A L CONTROL OF M A N N E D SPACE V E H I C L E S

If we now set the maximum permissible value of may examine the consequences when either C

111

= Cf m a x, then we

> Cu

(24)

> Qi/hV

(25)

w

or C

w

Clearly relation (24) presents no problem since it can always be controlled, and relation (25) becomes the governing factor. Thus, where Qi and V are known and Cimax is specified, we must have ki > Qi/Ct^V

(26)

to satisfy the condition, d < Cim&x. Relation (26) dictates broadly the condition of maximum safety if we set Qi at the value corresponding to the maximal production rate of which the biological system is capable. Alternatively, we may be forced by practical considerations to achieve a value of ki sufficient only to cause d to decay from transient peak values down to less than Cimax within an acceptable time-concentration envelope (Fig. 6 ) . Parametric description of the module involves estimation of the system variable, kiy in terms of the outputs and permissible concentra­ tions of the component gases. For design purposes, it is helpful to specify the parametric nature of ki by noting again the assumption that ki = rat/V

(27)

Substitution of this into the constant coefficient, Qt/kiV, of Eq. (23) gives kit

C*w = (Qi/rai)(l - e-™) + de~

(28)

which makes it evident that at t = oo the component concentration will be determined by the ratio of production rate to ventilation and clearance rate. The design specifications for ventilation, r, and absorbing factor, ai} are therefore set by the most effective combination of these respective parameters compatible with the physiological optima for d in the steady state. For a given permissible concentration, we must select values of r and at such that Qi/Ci = rai. In practice, solutions require empirical evaluation in terms of the physical capabilities of the circulation pump and physico-chemical properties of the absorbent or regenerative system for the particular (ith) gas under consideration. It is equally easy to see that the overall pumping rate, r, is deter­ mined by: (1) the C i m ax of that gas showing at once the highest toxicity and lowest removal absorption affinities, or (2) the gas showing maximal production rate, or (3^ both. In Fig. 7 the slope, Qi/Ctai = r, denotes

112

ROBERT E. SMITH

schematically the minimal pumping rates compatible with steady state control of a system containing a gas of volume production rates, Qi/Ci} for given values of a*. The presence of another gas competing with the ith gas for the absorption or scrubbing mechanism would have the effect of reducing at as would the progressive poisoning or depletion of the scrubbing system. While in this most elementary theory (cf. Fig. 7) any reduction in .

r=40

r=20

0

FIG. 7. Graphical illustration of steady states in which Qi/d = rai (see text); units arbitrary, but dimensions of r must be the same as those of (Qi/Ct), e.g., liters/min or mol/min.

adsorbent efficiency would appear to require compensatory increases in pumping rates, it is quite evident that in practice the breakthrough characteristics of adsorbents would largely preclude such a solution. In effect, therefore, we are assuming that a* does not become limiting within the physiological range of gas concentrations encountered. Essential considerations in the design of the scrubbing and regener­ ation system would include deciding the most efficient geometric arrange­ ment of pumps and regeneration units. Evidently a series arrangement would permit minimal pumping volumes, but it would also be subject

E N V I R O N M E N T A L CONTROL OF M A N N E D SPACE V E H I C L E S

113

to the limitations associated with series resistances and with differential rates of degradation in the respective regenerating components, e.g., those for C 0 2 , H 2 0 , and noxious vapors of various origins. The least effective absorber would dictate the over-all flow rate whereas with independent parallel systems, the total flow would be the sum of the respective r values required for each component. This would have the advantage of independent control both in design and operation, but probably at a higher cost in payload. Specifications of net volume of the module demand an optimized solu­ tion to the combined requirements of physiological maintenance. This is done with respect to gas exchange and thermal regulation, together with the operational space requirements designated by human factors analyses and the objectives of the mission. Here are considered only the first two, for these, in effect, are primary in determining minimal modular dimen­ sions. Multiplying these dimensions to provide larger total volumes re­ quires, in general, certain corrections of a non-linear nature because of accompanying changes in wall surface relative to volume capacity. Thus, the thermal outboard loss rate is proportional to the square of the linear dimension, while the heat production within the vehicle will tend to vary as the cube of the linear dimension. The resultant leads to the wellknown f-power law which requires the surface of bodies of similar shape to vary as the f-power of the volume. In the present application this means that if we double the volume, the surface will increase by a factor of only 1.59, or surface/volume will be reduced from an arbitrary value of unity to 80 per cent of the original ratio. Therefore, if we were to build the two-man Gemini satellite from a similar design, but with twice the volume of the one-man Mercury cap­ sule, our surface would have increased by a factor of only 1.59 times that of the original Mercury model. Hence, for the same environmental tem­ perature, the heat losses via the surface would have to be increased by a factor of 2/1.59 = 1.26 per unit area, i.e., by the cube root of the volume. Stated otherwise, as we increase the number of modular units the surface heat losses per unit of volume will become less; correspondingly, how­ ever, thermal balance requirements will impose in non-linear proportion greater demands upon regulatory heat exchangers. One interesting, if somewhat academic, corollary to the heat balance of these vehicles arises from the fact that heat production of mammals (and other animals as well) does not increase exactly as predicted by the f-power law; instead, heat production is proportional to body mass raised to the 0.73 power [39]. This can only be possible for bodies which undergo systematic changes in shape as a function of increasing body mass (volume). Therefore, it can be suggested that for a constant pro-

114

ROBERT E. SMITH

portion of wall losses to biologically occupied volume, an optimal solution would be to change the shape of the compartment in a systematic way that will cause the external surface to increase as the 0.73 power of total volume. One of the ways of doing this is to dispose the volumes into a series of double cones in which the ratios of base to altitude are varied systematically [40] such that the cones become relatively slimmer, i.e., of higher altitude vs base diameter, as total volume is increased. By introducing a changing shape factor of this nature into the design of the compartments, one could, at least in theory, retain complete linearity between respiratory and thermal regeneration capacities and the numbers 10,000

9000

8000

7000

1

2

3

4

5

6

l / k i( m i n ) FIG. 8 (see facing page for legend).

7

8

ENVIRONMENTAL

CONTROL OF M A N N E D SPACE V E H I C L E S

115

of human modular units. In lieu of this it is obviously necessary to as­ sume rising capacities for heat exchangers as the number of commonly connected modules is increased. As noted, this will increase approximately as the cube root of the total volumes. The net volume of the compartment may be specified in terms of the relation (27) in which hi — rai/V, but with rat now established by re­ quirements of steady state maintenance. Thus the volume, V, may be explicitly defined by selecting a value for ki consistent with physiological restitution of the limiting gas concentration. Assuming, therefore, that Ciy the concentration of the ith gas (e.g., C 0 2 ) must fall from 1 to 0.03 in 5 half-times and that we would allow 10 min for this to occur, the 1 half-time would be 2 min; whence thefc«= 0.693/2 = 0.3465 m i n " . The volume is thereby defined by V = ra^/0.3465, or one may arbitrarily graph curves of V versus 1/ki in which the slopes indicate the required values for rat as illustrated by Fig. 8. Thus, for a given half-time, the volume may be increased only by a proportionate increase in the slope, through appropriate alteration of the pumping rate, r, or the absorption coefficient, aiy or both. An alternate computation for V is obtained by recalling from Eq. (23) that at t = oo V = Qi/Cjci

(29

which differs from Eq. (27) only in a literal sense, since rai = Qi/Ci. Hence, both data appear in the legend of Fig. 8. FIG. 8. Volume, V, of chamber versus half-time of clearance for various steadystate concentrations, d (in per cent at PB = 760 mm Hg), and rates of production of Qi (in liters per minute STP) of the ith gas. The numerical values of concentra­ tions and production rates also correspond to real values which would apply if the ith gas were C 0 2. The slopes {V/hi) have the dimensions of liters per minute, but also may be represented by the values of rat in mols of the pure ith gas which must be removed per minute to maintain the respectively indicated values of Ci. Condi­ tions la, etc., are indicated in the accompanying legend.

Condition

d (%) (concentration) at t = OQ

Qi (liters/min) (Production rate)

la lb lc 2a 2b 2c 3

0.03 0.03 0.03 0.30 0.30 0.30 2.00

0.600 0.210 0.175 1.340 0.600 0.210 0.600

rai (mol/min) or (liters/min) (Slopes of V vs 1/ki) 90 32.8 25.6 20.0 9.0 3.3 1.4

2020 734 574 447 202 73 30

116

ROBERT E.

SMITH

These curves illustrate several simple facts concerning the turnover times for chambers of various volumes in relation to physiological limits. Thus, if we are to insist on very rapid clearance rates, e.g., halftimes of two minutes, the total volume cannot exceed three times the per minute pumping volume. The latter is itself specified by the ratio: metabolic output/concentration chosen. In general terms the lower the steady-state concentration required, the faster must be the pumping turnover rate, both absolutely and in relation to total volume. Thus, if we choose a modular volume of 2000 liters and a maximal C 0 2 output of 0.6 liters/min, a pumping rate of 202 liters/min, i.e., 10 per cent of total volume, will suffice to maintain the steady state C 0 2 concentration at 0.3 per cent, whereas 2020 liters/min (100 per cent of total volume) would be required to hold C 0 2 levels down to the sea level equivalent of 0.03 per cent. At the permissible standby concentration level of 2 per cent C 0 2 , a pumping rate of 30 liters/min (1.5 per cent of total volume) would accommodate maximum metabolic output. At this rate, a 50 per cent failure of the scrubbing system would allow the steadystate concentration to rise to only 4 per cent C 0 2 , which is still tolerable, though certainly inadvisable for extended periods. B. Carbon Dioxide Concentration

Data on C 0 2 concentrations and production rates within physiological ranges are summarized in Fig. 9 along with curves describing expected steady-state values for a series of pumping clearance rates ranging from 50 per cent to 1.25 per cent of total volume per minute. From this figure it is clear that to remain within the range of normal production rates and concentration levels, pumping rates will have to be of the order of five or six per cent of total system volume per minute. This is also illustrated in Fig. 10, where for various production rates the respective C 0 2 concentrations are plotted as a function of the turnover fraction of total volume required per minute to maintain these steady-state levels. Included for comparison are data obtained by Armstrong [41] in some excellent studies on ventilation requirements of men in closed compart­ ments at various C 0 2 production rates induced by graded work levels. These data correspond closely with present predictions regarding relative ventilation rates requisite to maintenance of given levels of C 0 2 concen­ tration. They further emphasize the possibility that the maximum Q Co 2 assumed in Fig. 10 would be too low for work exceeding the equivalent of a 2 miles/hr walking rate. The present estimate of five to six per cent per minute turnover would, however, permit an output of 1 liter/min C 0 2 at ambient concentration of one per cent or less, which appears adequate.

ENVIRONMENTAL

CONTROL OF M A N N E D SPACE

VEHICLES

117

5.0

4.0

3.0

2.0

1.0

e,5

Q co0

1 min

FIG. 9. Physiological tolerance levels of C 0 2 (in per cent) are indicated on the ordinate and the range of C 0 2 production (in liters/min) on the abscissa. The curves intersecting these coordinates predict the steady-state concentration for a given production rate or vice versa; the slopes are expressed as fractions of the total volume cleared of C 0 2 per minute, whence the lower the clearance fraction, the higher the slope and the greater the steady-state concentration for a given produc­ tion rate.

Of current interest is the ventilation ratio of the suit loop for the Mercury satellite system (Fig. 11) [27, 42] which approximates a five­ fold volume turnover per minute.* This ratio, if confined exclusively to the suit loop, is some 50 times higher than would be necessary for C 0 2 8

* Gas flow through the3 suit loop is stated [27] to be 10 ft /min, with enclosed net volume of a nominal 2 ft (Courtesy of Dr. James N. Waggoner, The Garrett Corp., AiResearch Mfg. Div., Los Angeles, California).

118

ROBERT E. SMITH

control alone. The flow system, however, was also in series with a heat exchanger rated at 1000 BTU/hr (~252 kcal/hr), which is about three times the standard human caloric output at rest. Notably this flow rate, if put through a suit channel, e.g., at torso level (assuming the gas space to be a 1-cm annular ring of inner radius 16 cm), would put the gas

C.%C0o l 2

3.0

jl1 \

1

11

1

1 1 1

t

tl 2.0

1

1

11 >l

' ! i

1

i •

\

i T

1.0

\ \

1

A ' A •I

\

\

v

\

\\ \

\

\ \

\

\

0

0.05

0.1

0.2 f V min"

1

0.33

0.4

0.5

FIG. 10. Steady-state concentration of C 0 2 obtainable at various fractional venti­ lation rates, /. Solid lines represent points from intersections on Fig. 9 for "minimal" Qeo2 (lower line) and "maximal" QCQ2 (upper line). The dotted lines are computed from data of Armstrong [41]; the upper curve for C02 production of 2.4 liters/min obtained with subjects on a 5 mi/hr work schedule, the middle for 2 mi/hr and the lower for bed rest.

velocity at 1.63 km/hr. Using the upper limit of temperature given (26°C) and PB = 250 mm Hg, one may estimate from Fig. 2 that the 2 mass velocity, G, was of the order of 600 kg/meter /hr, which is well within the region of stabile heat conductance (cf. Fig. 3). We may note also, however, that at PB = 760 mm Hg, or PB = 500 mm Hg, the same conductance would have been achieved at gas velocities of \ to \ of the values required in Mercury. This may have influenced the choice of a near sea level atmospheric pressure apparently employed by the Soviet

ENVIRONMENTAL

CONTROL OF M A N N E D SPACE V E H I C L E S

119

Vostok manned satellites. The composition of the atmosphere as actually employed, however, has evidently not been revealed beyond a statement by Parin* that it contained 20 per cent oxygen. As it appears that the mixture is not ordinary air, one may speculate that it may contain helium, perhaps in addition to nitrogen. Three advantages to this would be: (1) the conductance of helium, which is about 15 times that of nitrogen; (2) the advantage in mass accruing from the use of helium; and (3) protection against "bends" or aeroembolism in the event of acci-

FIG. 11. Schematic diagram of the Mercury environmental control system (from Johnston et al. [42]).

dental decompression, though in varying degrees, depending on total pressure and other conditions of the mixture. The chief disadvantage is the high diffusivity of helium and the consequent leakage risks entailed in its use. C. Oxygen Requirements

Oxygen (Qco 2/0.8), to the C 0 2 relations at

requirements correspond approximately with the value whence the program for normal usage can be linked directly sweep-out level as governed by the system constant. From t = °o when d = Qi/rai} we would have Qo* = Qco2/0.8

* V . V. Parin. Observations on the physiology of animals and man during space travel. Symposium on Soviet Space Biomedical Research, Univ. Calif., Los Angeles, California, May 8, 1962.

120

ROBERT E. SMITH

and Qo* = racot (Cco,/0.8) D. Atmospheric Regenerators

Gross approximations of maximal efficiencies of chemical regenerator systems may be computed upon the assumption that the indicated stoichi­ ometric reactions will go to completion (Table I ) . This, of course, ignores TABLE I .

COMPARATIVE THEORETICAL YIELDS FROM ATMOSPHERIC REGENERATION SYSTEMS Products fixed ( —) or released ( + ) in per cent initial reactor mass

Reactor system Liquid O2 Converter 0 2 Converter + LiCl—LiOH K02 K 0 2 + LiOH H 20 2 H 20 2 + LiCl—LiOH*c H.O2 + LiCl—LiOH

H 20

C02

o2

0 -18 -25 -25 +53 -18 +36

0 -45 -31 -36 0 -45 -44

+83 +27 +34 +33 +47 +15 +16

... . coefficient (Per cent) 83 90 90 93 100 78 96

0

Adding 0.1 mol LiOH/mol K 0 2 to give a CO2/O2 ratio = 0.8 (see text); the sum (93)6 is correct before rounding off constituents. Assuming 0 2 at 47 per cent yield from H 20 2 and that the H 20 yield from H 20 2 will be added directly to water supply. Initial reactor includes mass of one mol O2. "Including H2O2 as initial reactor and adding H 20 terms without respect to sign.

a host of well-known physical factors determining both relative com­ bining proportions as well as rates of reaction [cf. 24, 43, 44]. Thus, for the system using LiOH and LiCl and cryogenic oxygen at 83 per cent net yield (cf. [45, 46]) we could write the over-all reaction system as (LiOH + LiCl) + H 20 + C 0 2 + (0 2) = L i H C 0 3 + LiCl • H 20 + 0.83(O 2) + 0.17(00*

in which we may take the initial mass as the sum of gram equivalents of the bracketed terms. As a per cent of this, the respective yields of ab­ sorption are: for H 2 0 , 18 per cent; C 0 2 , 45 per cent; and release of 0 2 , 27 per cent; whence an over-all utilization coefficient would be the sum of these, or 90 per cent. These values may be compared with those obtainable from a process * Relative converter mass equivalent (cf. [45]) assuming 83 per cent converter efficiency.

ENVIRONMENTAL

CONTROL

OF M A N N E D

SPACE V E H I C L E S

121

originally developed by Reynolds [47] in which potassium superoxide serves for both C 0 2 and H 2 0 absorption and for 0 2 release [cf. 29, 44, 48]. The reactions of K 0 2 with either water or C 0 2 proceed spontane­ ously with the evolution of considerable heat. These may be written as follows: (1)

2 K0

2

+ H 20 = 2 K O H +

f

(2)

2 K0

2

+ C0

1 0

2

= K 2C 0 3 +

0

2

2

Reaction (1) evolves some 9.4 kcal/mol and (2), 43.1 kcal; the further reaction from the product of (1) would proceed in the presence of C 0 2 to 2 K O H + C 0 2 = K 2C 0 3 +

(3)

H 20

and finally, since the deliquescent K 2 C 0 3 is easily hydrated, water stor­ age would be facilitated by the reaction: (4)

4 H 2 0 + 2 K 2 C 0 3 = 2 ( K 2C 0 8 • 2 H 2 0 )

From the net reaction (5)

4 K0

2

+ 4 H 2 0 + 2 C 0 2 = 2 ( K 2C 0 3 • 2 H 2 0 ) + 3 0

2

the respective yields, in per cent of K 0 2 utilized, are: for H 2 0 , 25; C 0 2 , 31; and 0 2 , 34; giving a total utilization coefficient of 90 per cent. On this basis K 0 2 appears about as economical as does the 0 2-LiOH system. The latter, however, offers some advantage over K 0 2 in that the respira­ tory exchange quotient (R.Q.), i.e., the ratio of C 0 2 / 0 2 , can be compen­ sated merely by regulation of the 0 2 supply. With K 0 2 this is not the case, since this ratio is fixed at 0.66 [cf. reaction (5)], well below the normal biological range (0.74-1.0) and mean of greater than 0.80. The R.Q. is depressed to 0.74 in fasting, but to as low as 0.66 only in certain pathologies, in hiberation, and by experimentally induced ketogenesis. Feasibility of the latter two in space operations remains interesting though questionable pending further study. K 0 2 , apart from its relatively low mass and high utilization efficiency, offers certain other advantages in being a powerful oxidizing agent. As such it is capable of destroying a variety of volatile organic atmospheric contaminants, including those associated with body odors and flatus and in one closed compartment study appears to have reduced the aerial bacterial count. Practical applications of K 0 2 to submarine emergency atmospheric control have been successfully tested [49]. The compensatory adjustment for R.Q. required with K 0 2 may be developed essentially as follows: At the normal ratio of approximately 0.80, an estimated 603 liters/day of 0 2 (appropriate for a 154-lb "stand­ ard" man) would expend only 36 mol or 5.6 lb of K 0 2 / d a y , while the concurrent C 0 2 production would activate some 43 mols, or 6.7 lb of K 0 2 . Whence we would continually accumulate excess 0 2 at the rate of

122

ROBERT E. SMITH

0.8/0.66, i.e., 1.21 X 0 2 consumption. This would raise the total ambient a pressure and in all but a pure oxygen atmosphere enrich the po2 s a straight line function of 0 2 consumption and time. However, as the slope would change inversely with total volume, the effect would tend to a minimum in large compartments. Alternatively, one might reduce the output of oxygen relative to C 0 2 production by some means, e.g. (1) by storage of the excess oxygen either in a labile chemical form or by compression or liquefication; or (2) by diversion of a proportion, in the amount X, of the metabolic C 0 2 into another absorbent (e.g., LiOH) or possibly otherwise, but sufficiently to make the net ratio CQ 2 absorbed by KQ 2 + X 0 2 produced

=

Q

g

arbitrarily the respiratory exchange quotient of the biological system. Thus, for each mol C 0 2 reacting with K 0 2 , an additional 0.2 mol would have to be diverted into a nonregenerative fixation system. Judging from one account [28] it appears that the Soviet unmanned animal-carrying Sputnik 2 must have carried essentially this type of air regenerator since it employed "highly active chemical compounds" which "absorbed carbon dioxide and water vapor. . . and "yielded the neces­ sary amount of oxygen in return." The translated description states further that "duplicate miniature electric fans fitted to the purifier served to ventilate the cabin. The device was controlled by a sylphon pressure relay, which put out of action (italics mine) the most active part of the regenerating substance when the pressure in the cabin rose above 765 mm Hg." Presumably this amounted simply to a periodic switching off of the one fan driving air through the canister carrying the absorber-regenerator mixture while leaving in continuous action the parallel unit containing only LiOH or possibly LiOH and LiCl. Since the gross volume of the complete cabin assembly was about 25 liters and the dog, Laika, weighing 6 kg is shown surrounded by equip­ ment, etc., to at least two times her equivalent volume, the residual dead space must have been of the order of 8 to 10 liters. Assuming a 10-liter dead space and a 6-kg dog at an air temperature of 15°C, one may esti­ mate the on/off programming interval of the shunting device. Taking the 0 2 consumption as equivalent to 66 kcal/kg/24 hr [31], one obtains 55 ml/min 0 2 consumption at STP, and solving for X in (30), the C 0 2 to be absorbed by the LiOH canister would be X = 0.8(55) - 0.66(55) = 7.7 ml/min, the 0 2 equivalent of this being 7,7/0.8 = 9.6 ml/min.

ENVIRONMENTAL

CONTROL OF M A N N E D SPACE V E H I C L E S

123

For the stated pressure of 765 mm Hg at which switching to full shunt was activated, the 10-liter dead space would have acquired an extra 67.2 ml of 0 2 over that present at 760 mm Hg. The time required to accumulate this would then have been, 67.2/9.6 = 7 min, which is a minimum time interval as it arbitrarily assumed zero absorption of C 0 2 by the LiOH shunt during this period. The time period of full shunt "on" (or better, 0 2 generator "off") would be only the time required for the animal to consume 67 ml of 0 2 , i.e., about 1.2 min. With smaller relative dead space both intervals would decrease, and with optimally propor­ tioned, simultaneous flow through the two canisters, the shunting device would need to serve only as a vernier correction on the atmospheric composition. This would be essential, however, in order to compensate for physiological changes in the proportions of C 0 2 / 0 2 being metabo­ lized. Shternfeld [48], among others, has suggested that H 2 0 2 could be utilized for the 0 2 and water supply as well as for a source of heat. Since the decomposition of 2 mol of peroxide would furnish 2 mol of H 2 0 and 1 mol of 0 2 , a daily 0 2 requirement of 27 mol 0 2 would yield 970 gm of water, a good f of the usual daily intake. The minimal weight cost would be 1.83 kg, or 4 lb/day. Concurrent C 0 2 removal with LiOH would require 530 gm, or 1.16 lb. As computed in Table I, the latter system could prove among the most efficient, provided that the mass of the converter could be held low. The recent advent of the "molecular sieve" has made available a highly uniform, easily reversible adsorbent suitable for selective adsorp­ tion of a number of gases on the basis of their molecular sizes. Employ­ ment of these sieves for atmospheric regeneration in space cabins be­ comes especially attractive for manned space missions committed to operations of longer than about 18 man-days. This duration marks the crossover point at which the required mass of chemical absorbent begins to outweigh the minimal mass of the reversible molecular sieve. The latter adsorbs at a rate which increases as the partial pressure of the gas (e.g., C 0 2 or H 2 0 vapor) and can be purged either by outgassing to a vacuum or by heating and flushing with a purge gas such as nitrogen or cabin air. The theory and design applications of molecular sieve columns to environmental control systems are treated in an excellent study by Willard [26]. Combination of these sieves with a pre-stage of water adsorption by silica gel has been suggested, particularly since the latter does not require an intermittent cooling time after desorption by a hot gas purge [50]. To be noted in passing is the sensitivity of the molecular

124

ROBERT E.

SMITH

sieve to "poisoning" by water and especially by traces of organic vapors; hence in practice such sieves should be preceded by activated charcoal filters and/or some burnout device or oxidant [cf. 51]. The concepts of the paper by Willard are in a sense complimentary to those offered in the present report in that they treat respectively the two components of the regeneration constant as defined here by the identity hi = rai/V

(31)

Thus, where this chapter concentrates on the physiological factors that appear to specify the nature and variance in r/V, which is the relative pumping rate, Willard's report achieves a solution to the same problems by holding r constant and specifying variance in aiy the adsorption "coefficient." They are both, however, constrained to meet the air veloci­ ties and conductance requirements of the cabin environment as dictated by the physical and physiological elements of thermal regulation. E. Carbon Dioxide Reduction Systems

Reduction systems are a prime essential to life support, whether in space or elsewhere. Methods for achieving them are aimed at creating free oxygen through the splitting of C 0 2 . Basically, the simplest method is to induce thermal decomposition by heating the gas to around 2200°C with subsequent quenching and repeated recycling through the reactor. (This does not appear feasible for the space vehicle.) In addition, de­ composition at very low efficiencies can be effected by ionizing radiation [52, 53]. Possibly more feasible is the photochemical reduction of C 0 2 by a catalytic process employing ultraviolet light (cf. [54, 55]). Finally, there is the biological photochemical conversion (e.g. [45, 56-58]) of COs and H 2 0 into carbohydrate and oxygen. In this photosynthesis the energy derives from photons in the visible spectrum, and the reaction is rather an addition than a C 0 2 decomposition since the 0 2 is furnished from the water molecule (cf. [57]). The electrolytic reduction of C 0 2 may be achieved by a number of means involving reductive interactions with the alkali metals or their hydrides [52, 53, 56, 59, 60]. These processes all involve electrolytic cells employing mainly molten alkali carbonates or hydroxides variously mixed with chlorides to give eutectic mixtures of suitable melting points and conductance. Molten alkali carbonates under continuous electrolysis yield oxygen, with carbon as a by-product. For lithium the basic re­ actions are [59] LiC0 8 -> Li 20 + C + 0 2 L i 20 + C 0 2 -> Li 2C0 3

ENVIRONMENTAL

CONTROL OF M A N N E D SPACE V E H I C L E S

125

of which the second is facilitated by collecting C 0 2 in a molecular sieve. Shearer et al. [59] have pointed out that compared to sodium and potas­ sium, the high reducing potential of lithium metal particularly favors the elimination of oxides of carbon in the electrolytic product. Thus the cathodic reaction in which the metal appears proceeds to completion by the reaction 4 Li + L i 2C 0 3 -> C + 3 L i 20

while at the anode the carbonate yields oxygen and C 0 2 which in turn combines with L i 2 0 to re-form carbonates. These authors have obtained the best results in their prototype model with an eutectic mixture of 60 LiCl/40 L i 2 C 0 3 which melts at 506°C. Their electrolytic efficiency of 15 to 20 per cent was found to be im­ paired by the presence of water vapor in the input gas, which also re­ sulted in some hydrogen appearing in the output stream. Since the biological C 0 2 output is only about 83 per cent of 0 2 intake, it is neces­ sary to supplement the 0 2 supply by electrolysis of water in the amount of some 0.44 lb water per man-day. Del Duca et al. [56] have re-examined the use of LiOH for anodic regeneration of 0 2 from C 0 2 . The reactions are 2 LiH + C 0 2 - » LiOH + C 2 LiOH

electrolysis

400°C

> 2 LiH + 0 2

in which the major difficulty is the possibility of the formation of in­ soluble L i 2 0 by the reaction LiOH + Li

L i 20 + § H 2

The contents of the cathode chamber are pumped into a centrifugal separator and the LiOH and LiH are returned respectively to the cathode chamber and to the chemical chamber. Upon its return, the LiH reacts with cabin C 0 2 ; the resulting LiOH is returned to the electrolysis cell after being cleared of carbon. These authors also note the considerable weight advantage of this lithium system over that of sodium (275 lb vs 340 lb/man-day) as well as its much lower power requirement (1.4 kw vs 7.68 kw). The recent emergence of solid oxide electrolytes [60] appears poten­ tially to open a new approach to the problem of C 0 2 reduction. The disclosures [60] indicate that it has been possible to demonstrate C 0 2 reduction and 0 2 enrichment in small electrolytic test systems employing solid electrolyte discs made up of T h 0 2 - L a 2 0 3 and of T h 0 2 - Y 2 0 3 (others arbitrarily possible) under very close control of the lattice structure. The

126

ROBERT E. SMITH

theory postulates that "these solid solutions form anomalous mixed crystals in which there are vacant places (holes) distributed statistically at random throughout the anion component lattice. These make it pos­ 2 7 sible for 0 ~ ions to jump from 'hole to hole and, therefore, to migrate in an electric field." Hence, "based on the principle of electrical conduc­ 2 tion by migration of 0 ~ ions through the solid, an electrochemical cell can be constructed in which C 0 2 is fed in at one electrode and oxygen emerges at the other electrode, the whole system, both electrodes and solid electrolyte being held at a temperature in the range of 4 0 0 to 9 0 0 ° C with an electrical potential sufficient to cause the desired rate of decomposition of C 0 2 to occur." Regeneration of 0 2 by hydrogen reduction of C 0 2 appears the most promising of the techniques currently at hand [52, 54, 5 6 ] . There never­ theless is very little agreement about which of the several reactions in­ volved should be exploited. These reactions, according to Weller [ 5 5 ] are as follows: (1)

C02

+ H2

CO (discard)

+ ^

H 20 \^

H2 + (discard) (2)

(3)

C02

C02

+ 2H 2 1 I + 4H 2

>-

C (store)

+

1/2 0 2

2H 20 y 2H 2 + 0 2 y

CH4 C + (store)

2H 2 I

2 PL

In each of these the 0 2 is obtained by electrolysis of the water. Due to the respiratory quotient, C 0 2 / 0 2 = 0 . 8 3 , electrolysis of some further water [in ( 1 ) and ( 2 ) ~ 0.5 lb per day] will be required to make up the 0 2 deficit. It is pointed out [ 5 5 ] that both reactions ( 1 ) and ( 2 ) present serious disadvantages due to their very unfavorable equilibrium con­ stants, in consequence of which a fairly large yield of methane may be expected. In particular, ( 1 ) requires many recycles through the system with continuous condensation of water at each pass. Likewise ( 2 ) is "thermodynamically favorable only at a relatively low temperature" 2 2 (i.e., K = P H 2O / P C O 2P H 2 = 1.1 at 6 2 7 ° C and 4 6 at 4 2 7 ° C ) and also re-

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

127

quires many recycles even with the catalysts. This system is thus rated "feasible" but also suffers from "serious disadvantages." Scheme ( 3 ) , in which methane is to be formed by passing the C 0 2 — H 2 mixture over the catalyst at 150-400°C, is regarded as feasible. The 4 K (= PCH 4PH2O/PCO 2PH 2) is 7.4 X 1 0 at 327°C, at which temperature with C 0 2 : H 2 at 1:4 the yield of methane should be 9 3 per cent at equi­ librium. Moreover, the decomposition of C H 4 is thermodynamically 2 favorable at high temperatures (K [ = P H 2 / P C H 4 ] = 62.2 at 9 2 7 ° C and 3 9 1 at 1227°C). The suggestion is made that no catalyst be used in this carbon-producing step in order to facilitate the routine cleaning of the reactor. In line with these conclusions is the position taken by Acker and Stern [ 5 4 ] in accepting scheme ( 3 ) as the most feasible, but they evi­ dently do not plan to salvage the methane. Notably Foster and McNulty [ 6 1 ] , using scheme ( 2 ) , appear to be having a broad spectrum of C H 4 and CO contamination of the effluent and find also that these may con­ stitute from 3 0 to 9 8 per cent of the recycling load. Conversely, Ruderman [ 6 2 ] , using scheme ( 3 ) , reports essentially full conversion to meth­ ane and a 99.8 per cent expected yield during subsequent thermal de­ composition of methane in the range 2 4 0 - 4 0 0 ° C using a nickel catalyst on alumina. Though the calculations of Del Duca et al. [ 5 6 ] show a sub­ stantial free energy decrease (15.0 kcal) for scheme ( 2 ) reaction, it is evident from the temperature dependence of K [ 5 5 ] that the desired equilibrium cannot be expected at workable temperatures unless a suit­ able catalyst is forthcoming. F. Waste Disposal

Waste disposal aboard the manned spacecraft assumes importance in direct proportion to the duration of the mission and the size of the crew. Beyond the matter of health and personal hygiene lies the need for econ­ omy of storage and the reclamation of usable materials. In their extensive studies on storage, disposal, and reclamation of wastes, Zeff and his associates [ 6 3 ] have designated four classes, namely excreta, garbage, sewage (as wash water), and refuse. Of these the first three may be considered potentially available for biological recycling. A further subcategory, i.e., body surface detritus, appears in the form of hair, nails, desquamated epithelium, solids in sweat, saliva, and seba­ ceous secretions, which totals some 1 2 gm/day and contributes to the atmospheric pollution about 2 gm/day of solids [ 6 4 ] . On the basis of a detailed study, Mattoni and Sullivan [ 6 4 ] recom­ mend exclusive use of an unperfumed, nondisinfecting pure soap, prefer-

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ROBERT E. SMITH

ably sodium or potassium oleate or palmitate. Because no such soaps are currently sold, they will have to be specially made. They make the in­ teresting suggestion that bath water could be filtered and reused several times; or if suitably treated to precipitate the soap, it might even be consumed directly. Because the total residual volume would amount to some 2700 ml, its direct reuse would save a considerable cost in power. In addition, these authors have engineered a prototype central hygiene station employing a plastic overall type of bath suit. Finally, this paper is of interest for having undertaken an evaluation of microbial control requirements within the space vehicle. Regarding nutrient wastes, Zeff et al. [63] have systematically ex­ amined storage techniques for three classes of wastes, i.e., type A, from an open cycle with no wash water; type B, the same as "A" but with wash water; and type C , referring to a semiclosed or closed life support system. Their summary table for a two-man, three-day mission is in­ cluded here as Table II. It is concluded for the two-man, three-day misTABLE I I . MINIMUM MASS AND VOLUME OF ENVELOPE FOR CANDIDATE STORAGE TECHNIQUES SERVICING A TWO-MAN CREW ON A THREE-DAY MISSION (from Zeff et al [63]). Type A wastes Type B wastes Type C wastes Storage technique

Mass Volume Mass Volume Mass VolumeJ (kg) (liters) (kg) (liters) (kg) (liters) a

a

Limitations

Freeze drying

1.9

19

2.0

25

NA

Radiation cooling

0.4

32

0.4 +

43

0.3

29

Vehicle must maintain attitude control

Thermoelectric cooling

7.3

57

9.5

73

5.7

52

Requires auxiliary power

Solar heating

1.8

20

2.2

26

1.7

23

Vehicle must maintain attitude control

Electrical heating

5.1

50

6.9

64

3.1

49

Requires auxiliary power

Liquid disin­ fectant

3.2

29

3.3

38

NA°

27

Requires centrifugetype storage container and water in wastes

NA

Wastes must contain more than 16% water

° Not applicable.

sion that while the best technique would be storage by means of radiation cooling, the expected variance in attitude of flight would rule out this as well as the solar heating method in favor of the freeze-dry technique. The

ENVIRONMENTAL CONTROL OF MANNED SPACE VEHICLES

129

latter is also favored by the assumption that for the three-day mission there would be no reclamation of water from urine; therefore all waste water could be used for evaporative flash-off cooling. Where Type C wastes are involved, these could be packaged and stored in refrigerated food lockers as the latter became available. Disposal techniques for the longer term, accumulated solid waste residues embrace the requirement of ejection from the vehicle and in­ volve sterilization of the waste residue and pulverization or vaporization of these materials as a condition of safe and sanitary ejection overboard. Probably the best technique is to incinerate at temperatures in the range 750-1000°C [63]. Either solar or electric heat could be used. As long as the continuous wearing of pressurized or wholly enclosed spacesuits remains essential to our space missions, we will be obliged to develop special devices appropriate to the recurrent needs for voiding both urine and feces. The former appears reasonably well accommodated, for men at least, by the device described by Redden [65]; however, for the female astronaut some other design is indicated. Likewise, present spacesuits evidently make no provision at all for the voiding of fecal material. The processing and reclamation of potable water from urine, waste water, and condensates of evaporative body water have received a good deal of attention [25, 66-69] but no universally satisfactory process is currently available. In a fairly detailed summary of developments re­ ported at the end of 1961, Slonim et al. [69] reviewed progress in some twelve basic processes for water recovery, none of which were then op­ erational, although three or four in the developmental stages showed some promise of success. Among these are the vacuum distillation and pyrolysis techniques [70] and those designated respectively as vapor compres­ sion [25, 71, 72], freeze crystallization [73] and zone refining (cf. [74]). The Mercury satellite system is arranged to collect only the water vapor condensate from the cooling of the environmental air. The water is separated by means of a sponge from which it is periodically passed mechanically into a reservoir. Such recovery (about 6 lb each 24 hr) is planned as an emergency drinking supply, especially for use during the post-flight period [75]. Among the more recent reports is the study by Beutel [74] on water recovery by a reiterative freezing technique, a special case of zone re­ fining. It requires about the same energy as a double distillation, but has the advantage of being relatively free from zero gravity effects. A num­ ber of devices call for one or more adsorption or other expendable chem­ ical processes, which in lieu of regeneration immediately excludes their applicability to all but the missions of shorter duration.

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ROBERT E. SMITH

G. Systems Design—Comparative Aspects

Since atmospheric regenerators are an integral part of the life support system, they must also be optimized in terms of the over-all energy bal­ ance of the system. As the latter depends both upon the nature of the payload and the duration of the mission, considerable latitude may ex­ ist in achieving apparently satisfactory solutions to the regenerator com­ ponents. If we assume a given daily *net requirement of materials for life sup­ port in the space vehicle, the initial total mass of the system must in­ crease as a straight line function of the number of days involved in the mission. Thus, for the total initial weight of the system, W, we have W = Wo + mt

(32)

where W0 is the initial fixed mass of the system, m is the net increment in mass per man-day, and t is the number of days specified for the mis­ sion. The initial fixed mass, W0, is given by the relation between total power requirements (kw) and the power specific weight (lbs/kw). The curves of specific weight vs power are characteristic for the type of power specified. For reactors, specific weight decreases with higher require­ ments for power, while solar plants either show a somewhat flatter slope [56] or tend toward a constant at around 200 lb/kw [54]. Accordingly, in the power range below about 25 kw, the latter show weight penalties TABLE III. ATMOSPHERIC REGENERATION AND MAINTENANCE COSTS FOR FOUR SYSTEMS (Data from estimates of Acker and Stern [54]: Including assumptions of power supply at 200 lb/kw and use of molecular sieves). Take-off weights may be computed from Eq. (32). System Stored supply (liquid oxygen) ( 0 2 gas) Hydrogenation of C 0 2 (water) Photochemical C 0 2 reduction ( 0 2 gas) Closed biological algal conversion: (A) Fluorescent light (B) Incandescent light a

Wo (lb)

kw

IFo/kw

m (lb/day)

285 (20) 268 (3) 490 (0) 683 (3)

0.45

[633]

0.77

[636]

2.1

[325]

4.49« (2.84)b 6.25 (4.60) 1.94 (0.29) 3.05 (1.40)

5.04 7.8

[339] [256]

1708 1994

For 0 < t < 120 days; estimates of m include 1.65 lb food. h w = Wo + m(t - 120) for t > 120 days.

1.0 1.0

ENVIRONMENTAL

CONTROL OF MANNED SPACE V E H I C L E S

131

lower than those of reactors. Expected survival time of the mission (W — W0)/m = t, for a given environmental power plant obviously de­ creases where wholly nonregenerative support systems are employed, i.e., where m is largest, and conversely in the idealized solar powered, fully regenerative system, the slope, m, approaches zero. From the estimates of W0 and m in four schemes for atmospheric re­ generation with solar power (Table I I I ) , Acker and Stern [54] conclude that the most promising system among these is the one employing hydrogenation of C 0 2 combined with electrolysis of the resulting water and venting to space the C H 4 and other by-products including some C 0 2 . Despite the low per diem food cost of the photosynthetic algal system, its initial mass due to the weight of its subsystem power plant (5 to 8 kw) is so large that a crossover point with the C0 2-hydrogenation system is not reached short of 1000 days. Where regenerative life support systems are to be powered purely by chemical energy, certain additional constraints are obviously placed upon the choice of subsystems and fuels. For the latter, the standard free en­ ergies of known chemical systems do not exceed a value of 3.3 kw-hr/lb of reactants (cf. [45]). In essence, any regenerative subsystem yielding a pound of product at this same cost or more cannot compete with direct storage. From results of an exhaustive study by the TAPCO group [45], it appears that among the fuel-oxidant combinations feasible for life sup­ port, the most efficient is a propellant-atmosphere system employing cry­ ogenic H 2 and 0 2 (—AF = 1.66 kw-hr/lb) coupled with an hydroxy fuel cell. The resulting prototype is of itself an impressive example of a bal­ anced systems design which integrates optimized subsystems for cryo­ genic storage, air-conditioning, and the fuel cell power unit rated at 1 kw. As presented, the design (Fig. 12) will supply 1 kw of electric power and accommodate both ecological and metabolic requirements of one man for 3 days; however, design parameters and conditions are given for power levels up to 10 kw for periods of 1 to 20 days. This system is highly effi­ cient as a power supply and also biologically attractive in its use of non­ toxic reactants which give off both pure oxygen and water as by-products. Likewise, the design incorporates appropriate heat exchange between enthalpic components of the system, including the metabolic, to permit at the 1-kw use rate adequate cabin cooling and atmospheric dehumidification together with C 0 2 clearance through freezeout. An option is offered for cryogenic dehumidification of the cabin air, since C 0 2 re­ moval by LiOH would be cheaper in weight penalty during the first three to six days than the cryogenic freezeout system. The latter is competitive when the use rate is raised to 2.5 to 3.5 kw-hr, or a higher pCo2 accepted. Comparison of the TAPCO prototype with an operational design such

132

ROBERT E . S M I T H PROPELLANT - ATMOSPHERE SYSTEM SCHEMATIC

VENT OVERBOARD

TEMPERATURE INDICATOR CAPACITY METER