LEAP - A ONE-MAN LUNAR ESCAPE AMBULANCE PACK
Dennis So Carton^"
College of Aeronautics, Cranfield, England ABSTRACT
Seeks to anticipate an accident or medical emergency during the early phases of lunar exploration. The ambulance des- cribed l i f t s the patient alone from the surface and into
rendezvous with a waiting hospital ship in close orbite Such a mission l a r g e l y predetermines the minimum size of vehicle and i t s layout. LEAP i s i n consequence a constant attitude vehicle with a two position thrust vector produced by a rotating nozzle or combustion chamber. The influence of the vehicle scaling constants, of i n i t i a l acceleration and
rendezvous height upon vehicle size f o r both s o l i d and l i q u i d propellant propulsion systems are described.
INTRODUCTION
LEAP i s a one-man vehicle intended f o r operation from the lunar surface. The concept involves the launching of the LEAP payload into a trajectory that w i l l place i t close to an orbiting space station. A Ms o f t, f rendezvous w i l l permit the transfer of the payload into the space station. As foreseen at the moment, the orbiting space station w i l l be of an Apollo type, but with only two men on board. A "normal11 Apollo and three-man crew w i l l be on the lunar surface f o r an extended exploration ( l) o See P i g . l .
In the event of an emergency ( i . e . , the occurrence of an anticipated but nonpredictable serious accident or i l l n e s s ) , the affected crew member i s transferred to the waiting space station. This vehicle w i l l have a special emphasis on medicine and surgery and w i l l perform the hospital function i f i t s
Presented at the ARS Lunar Missions Meeting, Cleveland,
Ohio, July 17-19, 1962. This study has been supported entirely by LEAP'S own e f f l u x .
•^-Lecturer, Department of A i r c r a f t Propulsion.
^Numbers i n parentheses indicate References at end of paper©
D. S. CARTON
normal duties are interrupted© (The normal duties of a lunar orbiting space station during a period of manned exploration are seen as possibly navigational aids, communication l i n k s , advisory and command post, research control center, and ex- tended pl^ysiological zero 11 gw s t u d i e s . )
I f LEAP i s not made available at the time i n question (during the early lunar exploration phase), then alternative arrangements w i l l be required i n order to maintain the emer- gency service at the proposed level* The only possible a l t e r - native to something like LEAP w i l l require, i n the event of an emergency, the launching of the surfaced v e h i c l eο A rendezvous in close lunar o r b i t and transfer of the patient to the
hospital vehicle w i l l f o l l o w . I t appears certain that the necessity of supplying to the two Apollo Earth return vehicles the potential a b i l i t y to rendezvous and transfer the patient w i l l involve penalties in the sense of mass allocation and
also problems i n guidance and control much more severe and delicate than arise with LEAP.
This paper describes the salient features of a supposedly f i n a l i z e d complete LEAP system. Particular point i s made to emphasize the manner in which the definition of the mission and mission philosophy themselves direct the l o g i c a l build-up.
At a number of places, problems are considered i n some d e t a i l i n order to c l a r i f y decisions made with respect to operation.
I t represents a part of a study undertaken by the author upon the effect of f i n i t e propulsion systems and mission philosophy upon vehicles and payload©
MISSION
The LEAP mission i s to l i f t from the lunar surface a 240 lb payload (of one noncompos-mentis human plus equipment) and to place this in the v i c i n i t y of a waiting vehicle, which i s in a close lunar o r b i t .
The waiting hospital vehicle i s in a carefully controlled c i r c u l a r orbit at a height 2 χ 1θ5 f t (38 miles) above mean mare l e v e l . The LEAP vehicle possesses excess propellant and can d e l i v e r to the payload a velocity of 5400 f p s . This i s s l i g h t l y greater than c i r c u l a r v e l o c i t y . The excess, which i s 5$ i n energy terms, i s to permit the propulsive maneuvers f o r the rendezvous and docking procedures. No maneuver i s required from the waiting vehicle.
The LEAP propulsion requirement i s specified in terms of an energy height of 3 χ 10° f t , of which 93·3$ i s put into the payload i n the horizontal sense©
MISSION PHILOSOPHY
LEAP, with i t s human payload, i s considered as a l a s t ditch operation© There i s no secondary layer of safety processes and no abort machinery. I t f o l l o w s , therefore, that a very high r e l i a b i l i t y i s demanded of the complete system.
The o v e r - a l l philosophy has therefore been l ) wherever possible, to simplify; 2) wherever possible, to use one component instead of two, even at some mass s a c r i f i c e ; and 3) wherever p o s s i b l e , to place equipment in the hospital vehicle.
MISSION CONSTRAINTS : G-ENERAL
No controlling activity at a l l i s required, or needed, from the human payload. In f a c t , once the "countdown" has been i n i t i a t e d , the payload i s unable to interfere with the pro- cedure at a l l . Command problems require solution at this point. These are p a r t i c u l a r l y vexatious when the commander of the landed forces ends up as the payload.
Countdown i n i t i a t i o n , which can only be operated after pay- load has been secured, w i l l normally be operated by the personnel remaining behind.
Exceptionally the whole process of LEAP assembly, payload
"tucking-in", and countdown i n i t i a t i o n may be done by the payload. The process from tuck-in i s i r r e v e r s i b l e I
GUIDANCE
I t i s not possible f o r the i n i t i a l phase of LEAP launch to be observed closely by the waiting space vehicle. At the moment of ignition of LEAP, the distance between the vehicles exceeds 50 miles. In consequence, i t would not have been possible to use radio from LEAP to space station or the reverse as part of a take-off guidance system. Nor would i t have been possible f o r the space station to have detected the i n i t i a l LEAP trajectory by any means. In consequence, the i n i t i a l phase of the launch has had to be guided independently of either space station or lunar surface. In f a c t , LEAP uses an elementary two-gyroscope i n e r t i a l guidance f o r a large part of i t s f l i g h t . This involves a certain minimum of prelaunch setting-up in order to define the l o c a l v e r t i c a l and also the planned "turn down" direction.
With LEAP, a simple rigging system i s used. The v e r t i c a l
D. S. C A R T O N
sense i s determined by centralization of a plumb-bob. The turn-down direction i s obtained by sighting v i s i b l e land marks already located upon a lunar g r i d . The precise relationship between grid master line and hospital vehicle f l i g h t path w i l l be determined by the f l i g h t vehicle and remain a constant f o r a given phase of the exploration. RLnal offset angle w i l l depend on the perpendicular distance from f l i g h t path to launch point.
The f i n a l phase of the t r a j e c t o r y , leading to rendezvous and docking, cannot economically and r e l i a b l y be guided by an automatic sensing system within LEAP. In consequence, the f i n a l guidance i s commanded by the space station using visual techniques. (This procedure i s the inverse of that reported by Houbolt, Ref.2, of the work by Lineberry and Kizrbjun, fief.3·) Velocity and trajectory changes are made by LEAP to the external commands.
INITIATION OF LAUNCHING- SEQUENCE
I t i s essential that the l i f t o f f of LEAP occurs at an exact time relationship with respect to the orbiting vehicle. Since no communication can occur between vehicles at l i f t o f f or immediately before t h i s , synchronization must occur one r e v - olution of the orbiting vehicle p r i o r to l i f t o f f . With LEAP, the f i r s t step in the setting up sequence involves the erection of a very l i g h t a e r i a l . This picks up the "overhead" signal from the space station. A clock mechanism i s thereby t r i g g - ered which counts off one orbit to zero. I f the launch i n i t - i a t i o n circuit i s not completed, the clock mechanism w i l l be reset at each pass of the space station. I f the "launch"
c i r c u i t i s completed, then a countdown to i g n i t i o n and accur- ately timed l i f t o f f w i l l occur. Obviously, a simple time mechanism w i l l prevent l i f t o f f occurring i f insufficient time i s given f o r warm-up procedure.
TRAJECTORY"
Although the minimum energy trajectory to a given altitude involves a tangential departure from the launching point, LEAP takes off v e r t i c a l l y . This i s dictated from f a i r l y obvious aspects of safety. Obviously the "pure" tangential departure results in the launching vehicle remaining very close to the lunar surface f o r an unpleasantly long time. In consequence, even i f the launching took place from the top of a convenient r i s e , the vehicle would s t i l l only possess a few thousand feet of altitude at the end of burning. In addition the f l i g h t line would need careful selection so that the trajectory took the vehicle through valleys rather than through mountains.
In f a c t , the possible presence of a precipitous surface close to the launching site predetermines an i n i t i a l v e r t i c a l phase that i s maintained f o r several thousand f e e t .
HANDLING- ON THE LUNAR SURFACE
LEAP, enclosed within an environmental control capsule, i s carried at a fixed location within the lunar landing stage of the main Earth launching vehicle. Upon landing on the lunar
surface, i t i s kept within the "command module" with adequate environment control. I f a command post i s to be set up away from the command module, the LEAP w i l l be transferred as one of the essential items. Such a transfer w i l l normally be carried out with the LEAP components s t i l l contained within the environmental control capsule. This has a t o t a l mass of some 475 l b which, in the lunar environment, can be managed by two men. Nevertheless, should circumstances dictate, the sep- arate components of LEAP may be moved, erected, f i l l e d , and f i r e d by one man.
At the conclusion of an expedition, i f LEAP has not been used in the emergency sense, some 175 Earth-lb of seleno- l o g i c a l samples are transferred instead.
LEAP - DESCRIPTION AND OPERATION
There are only two methods by which a single main propulsion system operating continuously can be used to f u l f i l the r e - quirements of LEAP. The f i r s t of these uses fixed geometry and a planned rotation through the f l i g h t of both vehicle and thrust vector. Such a proceedings requiring a f a i r l y elabor- ate guidance system within the take-off vehicle. LEAP on the other hand maintains a constant attitude with respect to the lunar surface. The vehicle leaves the surface under v e r t i c a l acceleration only. This i s maintained u n t i l the energy i n the v e r t i c a l sense i s sufficient to carry the payload to the rendezvous a l t i t u d e . At this moment the thrust vector i s rotated through SO0 to produce the required horizontal accel- eration to o r b i t a l velocity. This i s achieved by the physical rotation of the nozzle in the solid propellant vehicle example, and by the rotation of the complete combustion chamber in the l i q u i d propellant vehicle* The reason f o r the selection of this arrangement i s due to the simplification of the guidance problem. The guidance requirement being the maintenance of the vehicle at a constant attitude : the rotation of the thrust l i n e occurring at a f i x e d time from l i f t o f f .
LEAP layout i s shown i n Fig.2A. Problems of dynamic i n - s t a b i l i t y are simplified, and attitude control f i r c e s mini-
D. S. CARTON
mised by disposing payload and two propellant spheres about the thrust line so that the center of mass coincides with the thrust line in both positions of nozzle or combustion chamber throughout the f i r i n g time. Neither the solid propellant swivelling nozzle a detail of which i s shown in KLg.2B (4)> nor the swivelling combustion chamber for the liquid propellant unit i s a severe engineering problem.
ANALYSIS
The question of selection of operating conditions to produce a minimum vehicle for the LEAP studies depends very much on the scaling parameters used (5)· In this paper i t has been assumed that the liquid propellant system i s working at a combustion pressure of 15 atm. resulting in an effective ex- haust velocity of 8000 fps. The solid propellant combustion pressure i s 68 atm. and the same effective exhaust velocity.
The effect of the size of scaling parameters upon the layload ratio are presented in î l g s . 3A and 3B. The values actually considered are, f o r the pressurized liquid propellant system Mt/iip = O.O5 and Me/mp = 2t5 and f o r the solid propellant system Mt/tep = 0.1 and Me/mp = 1·
The assumed distribution of mass of payload and other non scaling items i s given in Table 1. The results of the anal- y s i s are in Table 2. I n i t i a l accelerations of 5 and 6 lunar g
absolute (about 0.8 and 0.97 Earth gfs ) have been considered for the liquid and solid systems respectively. These are not quite optimum but have been selected in order to minimize acceleration on the patient. The change in acceleration throughout the operation i s not large as the overall mass ratio i s 2.4 in each case. The resulting a l l burnt acceler- ations are 12 and 14.4 lunar g absolute {about 1.9 and 2.3 Earth gfs ) .
The analysis so f a r discussed has been based on a rendezvous altitude of 2 χ 1θ5 f t . The effect of reducing this height upon the payload r a t i o f o r various system scaling constants i s presented in î l g . 4 . In fact the unmanned LEAP mass which i s 469 lb at 2 χ 1θ5 f t i s reduced to 435 l b at 1θ5 f t and to 384 l b at 2 χ 1(A f t .
Were the selection of altitude a free choice, then a con- siderable advantage could be obtained by significantly reduc- ing the rendezvous height. Unfortunately, this i s a problem involving the limitations of the orbiting vehicle guidance and control, and i n addition the possible hazard of some mountains of the order 3 x 3-0^" f t . In consequence, although orbiting heights of 5 χ 10^· may perhaps be possible i n the future, i t
i s thought the 2 χ 1θ5 represents a confidently achievable maximum f o r the described operations.
CONCLUSIONS
LEAP i s one of a number of possible solutions to the r e a l problem of making available a medical/surgical f a c i l i t y i n case of need. The problem w i l l be of greater magnitude during the immediate build-up period of lunar exploration. A deci- sion to provide LEAP as an exploration emergency ambulance service w i l l be possible without major changes i n designs already commenced. I t w i l l involve a nonreturnable mass of l e s s than 500 lb in the lunar landing vehicle.
The f i n a l l y achievable figure f o r LEAP mass i s very sensi- tive to rendezvous a l t i t u d e , to the propulsion system used, and to the design values achieved f o r the scaling constant of vehicle and propulsion. The mass i s l e s s sensitive to the vehicle i n i t i a l acceleration. Quite modest maximum acceler- ations may be used (to the benefit of the payloads health) without a large inherent effect on payload r a t i o . The d i f f e r - ence between achievable payload r a t i o with l i q u i d propellant and s o l i d propellant i s not l a r g e . The selection of a l i q u i d system i s based on i t s better behavior i n exposed lunar day- l i g h t ο
Table 1 LEAP payload : distribution of mass
Payload d e t a i l s
Man 175 lh
Pressurized stretcher plus breathing kit 25 lb Attachment structure 5 lh Power supply; communication pack 10 l b Vehicle d e t a i l s
G-uidance gyro platform 5 l b Control attitude 10 l b Control thrust vector rotation 10 l b Total 240 l b
D. S. CARTON
Table 2 LEAP details
Rendezvous height = 2 χ 1θ5 f t ; payload + disposable load
= 240 lb
Pressurized l i q u i d Solid propellant
0.05 = 0.1
Me/mp = 2·5 sec = L O sec
Χα = 5 g^ absolute = 6 g£ absolute
*α = 26 f t / s e c2 absolute = 3 1 · 1 5 f t / s e c2 absolute
= 0.373 0·358
% Λ α 0.582 = 0.581
Μα = 644 lb = 670 lb
%
= 374 lb 389 l b= 19 lb = 20 lb
F = I67OO poundal = 2086O poundal
= 520 l b f = 648 l b f
Ο
mP = 2.087 l b / s e c = 2.61 l b / s e c
= 5 . 1 lb = 2,6 lb
tb 179.8 sec = 152 sec
tb
5508 sec v e r t i c a l
Μα - man = 469 lb = 495 lb
= 76 Earth lb = 80 Earth lb
NOMENCLATURE
absolute acceleration, f t / s e c effective exhaust velocity, fps
acceleration due to l o c a l gravitational f i e l d , f t / s e c2 nondimensional r a t i o , 1 slug mass/l lb and 1 l b f / l pdl altitude above surface, f t
rendezvous a l t i t u d e , f t
energy height based on lunar surface, f t mass, lb
mass flow r a t e , lb/sec time, $ec
a = c = SL = Sc = h
hx = He = M
m = t
v s velocity, fps
vx = c i r c u l a r velocity at height hx X = absolute acceleration in gj, units Subscripts
α = start of burning e = engine
h = horizontal acceleration phase of trajectory i = end of v e r t i c a l acceleration phase
t = tanks
ν = v e r t i c a l acceleration phase of trajectory ω = end of burning
p s propellant L = payload REFERENCES
1 Paget, MoAo and Mathews, C.W0, "Manned lunar landing"
Aerospace Eng. 2 1 , 50 (1962).
2 Houbolt, J . C o , "Problems and p o t e n t i a l i t i e s of space rendezvous", Astronaut. Acta V I I , 406 (1961)·
3 Lineberry, E. C . and Kurbjun, M . C . , "Preliminary study of a manned control of the terminal phase of rendezvous using visual techniques," unpublished NASA Langley Research Center Report (February 2 1 , 1 9 6 l ) ; cited i n R e f . 2 .
4 Darwell, H.M., "Thrust-axis control i n s o l i d propellant rocket motors," Rocket Propulsion Technology, edited by D.S. Carton, W.R. Maxwell, and D. Hurden (Plenum Press, New York, I96I), V o l . 1 , p . l .
5 Carton, D . S . , "Minimum propulsion f o r soft-moon-landing of instruments," College of Aeronaut., Cranfield, England, Note 94 (July 1959); also Spaceflight Technology, edited by KoW. G-atland (Academic Press, London/New York, i960), p.325o
«4 DISTANCE FROM t=0 TO RENDEZVOUS 183 MILES ·» POINT OF CLOSE DIRECTION COINCIDENCE ORBIT OF VEHICLE MOTION
•
T• -
L t*0 1 CIRCULAR VELOCITY = 5400 ft/s«c^_,— --— \19 RENDEZVOUS 38 y' vG>f^ HEIGHT MILES LEAP / ^ 2x ΙΟ5 ft \ / c ° Fig. 1 Lunar escape ambulance and close orbit pickupD. S. CARTON
/^p- S\ V A SIDE VIEW WITH NOZZLE 2nd. POSITION / ) (O^ \ I
VERT,CAL CTAKE 0FF ) Τ \ /LA J ( I ACCELERATION POSITION, 1st. POSITION 12 3 4 5 6 7 β 1. PORT SIDE PROPELLANT SPHERE. \ S ^\ \ \ \ \/ \\ I \ PLAN VIEW WITH NOZZLE IN 2. STRONG FRAME LOCATING »· 7 \ \V \ \ / \ JL l^v 1 \ HORIZONTAL ACCELERATION 3. MOUNTING POINTS FOR GYRO I \R^>""J^ 1 \ LOCATION. V ^ 1 \ POSITION. 4. GAS TRANSFER TUBE. / jl^'j U \ Q~_ ~ I1 " "\ 5. STRONG FRAME UPPER PORT. -^f313^ \ \/' \ 1 \ 6. PORT ROTARY JOINT. ----- l| ) 7. PRESSURE CAPSULE / STRETCHER. ^- ' ', L-L /V Θ. PAYLOAD - MAN. PRESSURE / N. L^Z^ tt * SUIT NOT MANDATORY. / Figo 2a LEAP layout with solid propellant motor
POLISHED R OK IDE I 1 1 ^^^^ FACES
\
J GAS FROM COMBUSTION CHAMBER TO NOZZLE BOX • \ \////////////////////////////^¾ΤΤ^^/ ///y////////// / ë P.T.F.E. (FLUON)^^^^Ξvy^^^ ^ —
P-T.F.Z. (F LUON) INSERT SEALING RINGy
STEEL SNAP RING Fig# 2b Detail of rotary transfer tube of solid propellant layout D. S. CARTONRENDEZVOUS ALTITUDE 2 χ ΙΟ1 ft C = βΟΟΟ ft/see RENDEZVOUS ALTITUDE 2 χ ΙΟ5 ft Mc Mt Hc-3xl04ft. m#P MP C= ΘΟΟΟ ft/sec. 042 · " 0 ° ^ He = 3 xio'ft. ^ ο
oo i
mo< Mc M / ^—' mp M / I 001 1 " ^ " ~—•— IO OO//V^ -
α40'
2 5 0 01 ^ / £ ^^^^^^ |2'5 OO < / \ \ 5 OOI g 0-36-j i ,
1 >ί X. Q 035- IOO OO σ37' \ 1
033- /" SO Ol \ 0-32 ' °36' \ 03I- \ 030 " IO Ol 0-35- ^ IO OOI 0-29 • - I I I I I I I I I I | ° 1 2 3 4 5 6 7 8 9 «° 01 23456789 IO INITIAL ABSOLUTE ACCELERATION LUNAR g*s (x<*) INITIAL ABSOLUTE ACCELERATION LUNAR g*s (x <x) Figo 3a Effect of scaling parameters Fig. 3b Effect of scaling parametersD. S. CARTON
H = 3 x lO'ft. C = ΘΟΟΟ ft/sec Ml
INITIAL A B S O L U T E ACCELERATION LUNAR g* s (Vcx)
Figo 4 Effect of rendezvous height