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Considerations on vehicle design criteria for space tourism
Acta Astronautica Vol. 37, pp. 223-237, 1995 Copyright @ 1995 Elsevier Science Ltd
Printed in Great Britain. All rights reserved
0094-5765(95)OOOgg-7
0094-5765195
CONSIDERATIONS ON VEHICLE DESIGN FOR SPACE TOURISM Kohki ISOZAKI,
Akira TANIUCHI and Koichi
$9.50+0.00
CRITERIA
YONEMOTO
KawasakiHeavy Induwies, Ltd. i Kawa.wki-cho,Kakand@hara-shi, Gijk 504, Jqxm
Hirosbige
KIKUKAWA and Tomoko MARUYAMA
Fuji Heavy Industries,Ltd. l-1-11 Yonan, Utsnnonajwshi, Tochigi 320, Jqm
The trauspoxtation research committee of JRS (Japaaese Rocket Society) has begun conceptual design of vertical takeoff and landing fully reusable SST0 (Single Stage to Orbit) rocke$ type vehicle as a smndard vehiile modelfffsplzemurian.Tlledesigncriteriaofthevehicle have paid most attention to the reqt&.ments of service to meet space tour amusement. Thestandard vehicle, which has 22m body length and weighs about 550 tons at takeoff, can provide attractive tours of 24 hours maximum for 50 passengers into the low earth orbit with a variety of w flight pleasures such as experience of weightlessness and earth sightseeing. Within the reach of our near future rocket technology, the design utilizes MMC, CF/Epy and Tiiw advanad matuials. The twelve LOX/LX2 engines consist of two nozzle types, which can be throttled ad gimbaied during the whole mission time, perform v&Cal launch and tail-fust reentry to final landing associated with raodynamic control of body flaps within tolemble accelaation acting on passengers.
The purpose of the project is to give a new, ad commercialization of space strong motivation for transportation by the space tourism studying from very different point of disciplines, such as space medicine, enterprise or tour service, which have never been learned in the history of Japanese rocket development (Elgure 1). Thus the aerospace industries are expected to make considerable efforts for producing low-cost spece transportation vehicles by the ground of seeking such a new space transportation market, namely space tour rref.21. [Jqnl”~eQ Rode1 society> Rsseorch
Grovps
Figure 1. Research Disciplines of Space Tourism The research project of ‘Space Tourism‘ has been adDpad for future sp&zeactivities by the Japanese Rocket Society (JRS). The study guideline was fast clarified by JRS Committee for Academic Activities at the annual meeting in April 1993. where experts in various disciplines wae invited to hold a pane1 discussion on this new theme [re.f.l]. The panelists were Mitarai Professor of Aerospace Physiology at Nagoya University, Kyotani Chairman of TFCNOVA. ‘R&ai Managing Director of JADC! (Japan aircraft Development Cotporation) and journalist Akiyama who experiured his space flight by Russian Soyu&Iir as the first Japanese astronaut. Copyright@ 1994 by the Intemational Astronautical F?edemtion.All right med.
223
The transportation research committee, the members of which are assigned from major Japanese BQOSPBC~ industries, airlines and corporate members of JRS interested in this project, was organized as a workiig group studying the definition of design criteria, the puspective of operational flexibility and finally a sladard vehicle model in cooperation with the other three rcscBcch activities such as medicine, enterp~se and tour service of spaze tourism (Figure 2). This paper summarizes the first year status repOn Of the research committee activity [ref.31including the issues of propulsion system design tref.41. liquid hydrogen technology rref.51.
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values of which may be completely different from current popular a&aft tour or virtual reality amusement
JRS Transportation Research Committee _%?f%ors. -Kawasaki
HeavyIndustries, Od.
21.
--
Mimbishi HeavyM&es.
-
Fuji HeavyIndustries,Ltd.
-
lshikawajimaH&ma HeavyIndutries, bd. Motor Co., Ud.
-Nissan -
AliNippon AirwaysCo., .bd.
-
Teisan KK.
Figure 2. Corporate Members of JRS Transportation Research Committee
.
.
2. Cuukbne
of %.w
Soace
Tour Model
Ltd,
The transportation research committee discussed two space tour models including amusement and available cabin services as shown in Table I. The flight plan is scheduled to take&f from a spaceport in northern Japan. The t~o-circles orbital flight course will provide a variety of m time, which offas Pacific Ocean / South Am&a sightseeing course at daytime depamue. Aliica / Asia sightseeing course at night depararre and so on.
Tourism
Table 1. Tour Course and Expected Service
Based on the guideline fist proposed by JRS, a unique study on the services expected for the first phase of space tourism [ref.@ will help u) clarify the features of space travel concerning tour course and cabin services similar to those of commercial airlines.
U-Hours Orbital Flight
BDqtime Depamtre * Pacific Ocean
. soulh America
Sight-
n Night
seeing
The key consideration on the vehicle design criteria for space tourism is how the ‘Spaceline’ vehicle can realize varieties of space flight joy that are most expected tour participants. A suggestion given by Akiyama was that the commercial values of the space tour = fust the experience of weightlessness and second the earth sightseeing [ref. 71.
~o-circlcs Orbital Flight
Departure . Africa . Asia
_ Amusement
Longitude
Depends on Departure Tie)
Weightlessness Astronautical Observation Telecommonication
_ Cabin Service
Recognition Souvenir
_
[ deg
Brcakfut/Lunch /Dinner
Soft Drink
Others The vehicle design pvided for space tourism should take these kinds of service scenario into consideration, the
mAI Amund the Eartl (DayandNight Scene
]
Figure 3. Two-&&s
Orbital Flight Tour
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As can be seen by the ground track of Figure 3. the vehicle leaving Japan flies first over the Pacific Ocean to Argentina. Then it passes the center of Africa to cross India and China. The passengers can enjoy spectacular scenery of their home town from the altitude of 200km. At the second mund of orbital flight the vehicle flies over Hawaii. The vehicle &orbits at At&a to pass the reentry inlerface at an altitude of KMkmin the middleof China. 2.2 Vehicle Design
Factors
Discussing the space tour in the low earth orbit as above is helpful to provide good engineering imaginations to designers. The design factors of vehicle for space tourism are C as follows.
cubin ArranPemetU The most important but difficult subjects that the vehicle designer should take into consideration ae passenger windows for sightseeing and a room for experiencing weightlessness. Discriminating the real sightseeing from the computer play like virtual reality, the vehicle should provide the windows as many as possible (Figure 4). The medical research group suggested that some passengers would have to observe the earth sitting on their own seats by physical reason, space sickness [ref.8].
Altitude Akiyama’s report accotding to his exciting joy flown by the Russian space station Mir is meaningful. He commented that the higher orbit altitude does not necemarhy deserve to raise the value of space tour service [ref.7]. Thus the ahitude of 200km is assumed appropriate for the sightseeing orbit. Orbit Inclination Since the earth’s equatorial region is sparsely populated, as we know, most of the spaceports will be constructed at the northern hemisphere at relative high latitude rref.61. The locality of population is one of the major factors of business to determine the orbit inclination. Although launching from higher latitude has disadvantage in payload injection capability, it provides passengers attractive earth sceneries that cover the equatorial zone to the both poles.
Figure 4. Window Arrangement A microgravity amusement space will be necessary for passengers who will enjoy floating in microgravity condition preventing accidents from kicking other’s head
Tour Time The maximum flight time of 24 hours can be preliminary determined by the tour requirement to cover whole sight of the earth in the daytime. This maximum time is also a trajectory restriction returning to the home spaceports located at relatively high latitude. On the other hand, the minimum flight time that the transportation reaearch committee has planed is two circles around the earth. This minimum time came tiom the necessity of preparine deorbit chance. Since one third of each orbital time is in d&mess. the m time has significant meaning for selection of sightseeing course.
Figure 5. Microgravity Amusing The idea of facilities currently used by airlines such as galley, toilet, miscellaneous utilities. television screen for monitoring scenery and sightseeing commentary system will be helpful in designing cabin arrangement
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. .
al Restrrctrons The passengers suffer most critical acceleration at ascent phase to orbit. As is already recommended, the vehicle shouldn’t exceed the acceleration limit of 4 G from head to foot and 2 G in the reverse way tref.91. Taking account of the medical endurable load and its direction in mind, more relaxed accelemtion level should be con&&d to prevent passengers from blackout.
Another medical problem is space sickness. Many passengersare thought to suffer severe space sickness as soon as they areput in weightlessness. Although there are medical treatments such as oral or transdermal medicine. mechanical methods like LBNP (Lower Body Negative Pressure) or Seat head rest (Figure 6) would be also effective as the first aid [ref.8]. Head Rest /
\
=P
, which have smaller cross range capability than the nose-first ones, don’t require high attitude rotation maneuver. Giving the design feasibility for passenger priority to accommodation, the uansportation research committee adopts the tail-first reentry style. 3.2 Performance
Reauirement
The total velocity increment required by orbital conditions is summarized in Table 2. The orbitai altitude is 2OOkm with the inclination of 45 degrees. Velocity increment of 300m/s is assumed for the vertical landing maneuver and 290m/s is added to the total velocity as the performance margin. The resultant total velocity of 9.93km/s is relative large compared with that required for BETAbef.141 or Phoenix[ref.lS]. Table 2. Total Velocity Increment
Figure6. PassengerSeat Design Mission 3.
Phase Velocity Increment
Standard Vehicle Orbital Velocity Gravity Loss Drag Loss Maneuver Loss Deorbit Impulsive Velocity Descent Landing Maneuver ____________________--Design Margin Aseem
Although winged launch systems are attractive ad mentally me in expectationfor futurecommercial space transportationstyle on the analogy of current &l&s, vuticai takeoff and landingfuily reusableSST0 (Single Stage to Orbit) mcke4 type vehiiles are generally thought feaaiblc within near-term technology [nf.lOl [rcf.ll]. Thus the muqwution nseareh cunmittee adqwdthis
vutical mkeoffandlandingSSTO
thestmtddvehiclemodelforthepwentsp8cetourism study.
Total Velocity
&m/s1 7.701 0.900 0.600 0.070 0.070 0.300 _____0.290 9.930
rocket=
2.3 Confiwration
Design
Since he design criteria of the vehicle for space tourism dependmuch on the user’srequestof sightseeing and weightlessness experience. the cabin arrangemauhas significant influence on the vehicle’s configuration. The vehicle design results in the body length of 22m with the bottom diameter of 18m as shown in Figure 7.
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22m
k-----4
Figure 7. SST0 Vehicle for Space Tourism
lgm------/
A two-floors cabin equipped with many sightseeing windows and a microgravity amusement space makes the top of the vehicle in upright position. The propellant tank is a semi-integral structure using advanced material to reduce weight, which has common bulk head between LH2 and LOX propellant tanks. livelve e@nes. which consist of four booster ad eight sustainer engines with conventional bell nozzle. ae mounted in a cituular position around the lower tank structure. Noxakexpansion ratios are 15 for the bwster engines and 40 to 80 (two positions) for the sustah~~ eWmes.nspectivdy. The vehicle has four l&age type landing gears, which can be -ted in the body. The energy absorption concept is conventional oleo pneumatic system. Ib prevent topple &rum of the vehicle in case of cue kmding gear failure, the length of oleo stroke in the opposite position will be s&x&ted. Delta Clipper has pvsed an idea to extuxl the noxxle of a sustainer engine instead of tbehtikdlandinggear[ref.l6].
3.4 Cabin Figure 8 shows the comparison of two typos of cabin arrangement. One lines up passenger seats straight like current airlines and the other adupts the seats lined up in circle to provide better view through windows. Akiyama arrsnganentincheleasthebu&ine -theseat [ref.l’ll. An image of cabin amtngaWtwith~gbtlinedup seatsisreahzedbyascalemodelsbmvninFii9.Tbe microgravity amusement space surrounded by a tmnqamntwallwithanentrancedoor,whi&~totbe upperdeck,locatesattbenarpositionoftheubinflooc l!vo waste management compartments ain seen in the [email protected] flightofstairstotheupperdeckisalsore&ed. lbereamtwopilotseatsatthefmntsideoftheupper deck, from which enough wide view is available chrting reentry and landing operation (Figme 10). Aa airlo& seen in Figure 1 I is also equipped on the upper deck m emergency exit.
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Entrance
Cockpit
Door
Alflock
Amusement Space
UPPER DECK
In-line Seat Amneement
a: cockpit
PIlot Seat
Galley
Alh& /
Micmgravity Amusement space
8. Cabin Design
Stewardess
Seal
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Mass Charactcristic$
The mass characteristics of the present vehicle estimated by the first design work is shown in Table 3. Table 3. Mass Characteristics Subsystems 1. E%ucmrc ~Engincskirt
9Engine lluust smrc * Maincabiisaucture - &ckpitK.abii Smtcrure
2.
3.
4. 5. 6. 7. 8.
* ‘Ihamal Pm&don System Heat Shield (Including Sustainer EIG Cover) * Landing Gear System * Airlock System Prupulsion System * Sustainer Engine * Booster Engine * LH2-Tank * LOX-Tank . AuxiliaryTanks (LH2,LoX) . FkwuizationSystem (AHe and CHe) . Reaction Control System . Feed System . Residual Prolxllant (Excluded in Subtotal) Actuatca System . Auxiliary Power Unit . h4isccUancous Actuators and Pumps Environmental Conuol System Power Supply System Navigation, Guidance and Control System Communication and Data Acquisition System Misccllanwus Equipment * Passenger Seats * Crw Seats * Galley * Toilets
* Misauancous
Subtotal kg] 15,959 3,860 2,063 2,009 2,436 3,944 1.375 272 29.5 17 9,456 4,184 8,888 4,089 459 1.950 400 100 2,475 820 220 600 1.100 580 305 431 1,542 1.127 62 122 130 100
Empty Mass 9. Passengers and Crew Weight * Fiity Passengers 3,750 * crw (FourPersons) 300 * Luggage 270 10. Propellants (Including Residual Propellant) 1Fuel (LHZ) 70.703 *oxid&r(L0X) 424.2 15 Total Lift-off Mass (Including Design Margin of 597kg)
SO.254 [kg1 4,320
494,918
550,089 [kg)
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4. S-cm 4.1
Acrodvnamic
Since the slush hydrogen will have unknown system complexity in the propellant feeding. the Norman cryogenic hydrogen became the baseline. The major disadvantages using plug nozzle are considered the lack of roll control force and the inflexibility of attitude control for engine-out contingency. Thus the conventional bell nozzle type with throttle seems feasible for the the st&ard vehicle.
Desim
Des&n
The standard vehicle employs four body flaps aI the engine skirt section, the two windward flaps of which activate as’ angle-of-attack trim control and the two leeward flaps are utilized for lateral stability. In order to evaluate the lift-t&rag characteristics and the trim capability, aerodynamic analysis is performed using the calculation model shown in the left hand side of Figure 12.
The transportation research committee has spend a long discus&u to configure booster and sustainez engine numbers, the cycleof which is finally &&Mto employ ‘augmented expander’. There are two reasons to decide the lotal engine number. The first reason is to keep Uuuat-toweight ratio at liftoff between 1.3 to 1.5 using 80% thrust level of LE-7 engine enhancing reliability and life cycle. The another reason is so-called intact abort criteria to maintain the thrust-to-weight ratio more than one in case of two engine failure [ref.lO].
As a result shown in Figure 12, the vehicle can attain the lift-to-drag ratio of 0.3 and 0.4 at the an&+f-attack 20 and 35 degrees respectively at hypersonic flight. Since the lifting body characteristics is about one third of the current performance of Space Shuttle and lower than the predicted Delta Clipper capability [ref.18], this type of tail-fust reentry vehicle may not have enough cross range modulation capability to return from high inclination orbits to contingency bases.
9.2 PrmUon
The expansion ratio for the booster engine was selected 15 to drive maximum thrust at liftoff. Due to flow separation limit of engine nozzle at sea level, on the other hand, the expansion ratio of 40 for the nzb~ted nozzle position of sustainer engine is feasible. The sustainer engine realizes Ihe expansion ratio of 80 at the full extended position within available space and weight. The propulsion group did enthusiastic study to find out optimal combination of the booster and the sustainer engines [ref.4]. The summary is shown in Table 4.
Design
The study on propulsion system has first started with the discussion whahex the standad vehicle adopts normal cryogenic hydmgen propellant rather than slush hydrogen, and whether it should employ conventional bell notie instead of plug nozzle type engines.
.s
B 4
..-
..___
g
‘-
J
9”
-a”
40 -30 -20 -10 Angle-of-Attack (deg)
0
10
0.04 E .s g 0.02 s z
0
E z0 -0.02 P E -0.04 0 ii 6.06 -60
Fiaurc 12. Lift-to-Draa..mand Trim _~ ___.__.____ ~~ Char~teristicn
-50
40 -30 -20 -10 Angle-of-Attack (deg)
0
10
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Tkble 4. Engine Configuration lI&otT Engine Characteristics
Note) + fiied expansion ratio of 40 Asthenumbtrofsusmineratginehuzases,theorbit injection cqability gets significant gain including weight penalty spend for the extension nozzle equipment. This payload gain is due to the long duration of sustainer engine operatitm with higher specific impulse. Since the vehicle’s base illt4 is restricted, there is a limitation of sustainer engine mrmber. The sUmlard vehicle adopts a combination of four booster engines and eight sustainer engines as the baseline (Figure 13). During terminal flight and laadiag phase the two booster engines ae Wivated while the other two rue put in its idle mods preparing for engine faihue. Suruiner mine
Acceleration of rotation required for the rwcticn control system is 1 deg/s*. which results in about 7OOON class thruster using GOX/GHZ propellant. The total impulse requirement of 4.2mNs was e&rated to satisfy rotation duty for the earth sightseeing. lhnk baffle, propellant feed system, pressurixation and vent system at alsoiXWKqWUystudied.
One of the critical design issues and requirements far realizing the SST0 vehicle for space tourism is the structural design. In order to minimize inert mass, various material candidates and structure system have been surveyed as shown in Figure 14. which wereoriginally studied for application of futute ae!rqXe plane [ref.191.
)66@3y
BOOSILX Engine
0
I
O
0
It is imbcated that the present technology hasalmost luIchedtothetargetmasslevelfortheprimarysuuctme, but more development effort is needed to realize the cryogenic tank structure.
Structure materials applied to the present staWd vehicle am seleaed among the liiht weight combinatiats of conventialal TPS (TbenMl FVoteuion System) and the advamM structure system mentioned above llccadins to the thermal requirement. The result of the Mine structure is shown in Tbble 5 and Figure 14. 18m _____I Figure 13. Engine Anangemalt at Bottom of vehiile
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Unit Mass
Unit Mass
ko/m’
0
kC/m 2
4 g s 1 _.__._-.._ ........_.. 500
lti
1SiX
0.
2ooo
Temerature (Kl
Figure 14. AdvancedSe~cture Survey ‘IWe 5. Tkmal Condition of Baseline Structure Smkxu-dl ComponcnrlArea
Surface Temperature
Heat Shield
1400-1100
MMC + SA/HC(TPS)
Engine Skin
1100-800
MMC + WHCIJ-PS) Al Alloy + Tiiw(-rPS)
Up to 800
CdbiiMOS.5
4.4 Nor-.
Material& Concept
(K)
CF/Epoxy/HC+
LH2-T?lIlk
ml
LOX-Tank
Up to 800
h0.K
-Metal Mm-ix
SA/HC
-Supa Alloy I Honeycomb
Composiu
Tiiw
Avionics of navigation, guidance and control system for the present SST0 vehicle has no aitical design issue. Attention will be paid for the onboard software design for both the ascent and reentry phase that calculates not only the optimal trajectory guidance and control but also manage various fail- modes as safe as possible [ref.u)].
Tiiw(-l-PS)
+ Insulation CF/Epoxy/HC+
-Titanium
CF/Epoxy/HC
Guidance and Control Svstem
+ Intcmal Insuladon
I Multi-
-Carbon Fitul
4.5
Insulation
Wall Epoxy /Honeycomb
ECLSS
ECLSS (Environmental Control and Life Support System) has significant meaning to provide many passengers comfortable space flight. A redundant system concept is presented in Figure 13.
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Figure 15. Structural Concept of the Standard Vehicle C__________________________._________-_~-------_______~
THC:Tempe-mdHumidirycOnd ARs:hmaphcrsRevipliudmspw .CO2Rmal .TKS C,mmhmnt Conml
Figure 16. ECLSS for the Standard Vehicle
4.6 Powcr Su~dv
.Cs:f-&~=-&w~Y ACM :Atmqhm
Comqc4ih
?.b”imtil
FDS : Fus Dction amiSuppmsion
1000.0T
Svstcm
The present SST0 vehicle will be equipped with two
kinds of power supply systems, which are an electrical power system and a hydraulic power system. The electrical power system is comprised of Fuel Cell Systems that use cryogenically stored oxygen ad hydrogen reactants, Reactant Storage/Distribution Systems and Electrical Control Units. The electrical power system produces all the electrical power requked by ECLSS, cabin services and other onboard avionics equipment for the entire flight duration.
z
5:o.o
Altitude [km]
500.0 z %
$0.0 0.0
Altitude [km]
Figure 17. Engine Char~teristics
I
30.0
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z
--
E -i+ F
---_
0.0
3
2
[s]
a
I
0.0
Tfne
0.0
Time [s]
350.0
‘Time [s]‘
, 350.0
350.0
Relative Velocity [km/s]
700.0 ?
0.0 8.0
1
i
/ 0.0 00 250.0
Relative Velocity [km/s]
T
00
Time [s]
1000.0
T
,-------r
.i 0.0
lime
[s]
0 Figure 18. Ascent Trajectoav
350.0
1000.0
f
I 00
0.0
:
:
;L
Time [s]
Time (s]
Figure 19. Reentry Trajectory
1000.0
350.0
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it possible to apply very light-weight material for protecting the thermal load on the base surface. 200
One of the critical performances for this kind of SST0 vehicle is the cross range capability. m calculation predicts that the cross range achieved by the aemdynamicbankmodulationwiU healitUelatgerthan 200km (Figure 20). Although the vehicle can not p&rut high maneuver in order to land any various m this will be huge enough to assure emegcllcy landing to contingency spaceports.
-S 5 100 8 4S
0
$ -100 0e -200 300
t -
’
’
’
’
2000
1000
’ k ’ ’ ’ *’ 4000 3000 Down Range [km] ’
*
’
’ ’ 5000
Figure 20. Range Modulation Capability For high intensity, but rather shorter duration power requirements that is called for by rocket engine gimbal actuation systems or by the body flap actuators, the hydraulic power system provides the necessary hydraulic power which is generated either by Auxiliary Power Units or by the rocket engine driven hydraulic pumps. In the case when the APUs are to be used, those will be opemted with the hydrogen/oxygen fuel to ensure the environmental friendliness. The APUs will be opemted only during the ascent and descent phase of the flight Operation.
5.
Traiectorv
Analvsis
As shown by the engine characteristics (Figure 17), the sustainer engines extend their nozzles to the seccnd position when the specific impulse of expansion ratio 80 exceeds that of 40 near the altitude of IOkm for ascent phase. A typical trajectory is simulated as shown in ‘Figure 18. The four booster engines begin throttling when the vehicle reaches the acceleration level of 3 G at about 100 seconds after liftoff. The vehicle achieves the altitude of 200km in 6 minutes. trv Tru
The guidancesimulating the reentry trajectory uses a conventional algorithm of drag acceleration control widely applied to winged vehicles. Since the acceleration acting on passeargus is limited 3 G. the nominal reentry load factor considering safety margin is designed 2.5. A typical reentry trajectory is shown in Figure 19. The low heat input due to the relative small ballistic coeffiiient and a large radius of the base curvature makes
. 6. Coneludlnn Studies on SST0 vehicle for commercialization of space tourism is about to start. The mn research committee was organized to define a sgldsld vehicle that is expected to be a real model for the space tour flight. It held nine meetings in total discussing on the vehicle design for the first yea. The committee has finally build a l/20 scale mockup shown in Figure 21. There have been many issues newly found md discussed at every meeting. Wb have rea@zed that are of the most critical technology challenges is mataiak Although the standard vehicle adopts various adnrred light-weight materials as much as possible. we had to accept 15% mass reduction horn the estimation hased on the cunent technology. Wegraduallyunderstandthatwehavetodumpourold fashioned way of thinking. Wb will conclude the present paper with the question “Can the enthuaiaam for Space Tourism save Japanese ‘stereotyped’aem+g~~ induauies ?” Probably it will be answemd by “Yes” in the nex future.
The authors are grateful to the members of the JRS Transportation Research Committee for space tourism, and especially Prof. M. Nagatomo and Mt. Y. Naruo of the Institute of Space and Astronautical Science, Mr. T. Ruikai of Japan Aircraft Development Agency and IL: P. Collins of National Aerospace Laboratory for their kind advises and useful discussions.
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Figure 21. Mockup of SST0 Vehicle for Space Tourism
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Referentas
(11 Nagatomo, M., “On JRS Space Tourism Study Program,” the Journal of Space Technology and Science, Vol.9. No.1. 1993. PI Collins, P., “‘IbwatdsCommercial Space Travel,” the Journal of Space Rchnology and Science. Vo1.9, No.1, 1993. r31 Isozaki, K., lhniuchi. A., Yonemoto, K., Kikukawa, H., Matuyama T.. Asai, T. and Murakami. K., “Vehicle Design for Space Tourism,” the 19th International Symposium on Space Technology ad Science, ISTS 94-g-22p, Yokohama Japan. [41 Mori, K.. Suzuki, A.. Iihara, S. and Nakai, S.. “Requirements on Propulsion System Design and Operation for Space Tourist Car&r Vehicles,” the 19th International Symposium on Space Technology sod Science, ISTS 94-g-23p. Yokohama Japan. El Hanad&T., Nagatomo, M. sod Naruo, Y., “Liquid Hydrogen Indusuy : A Key for Space ‘Iburism,” the 19th Intanational Symposium on Space Technology and Science, ISTS 94-g-24p. Yokohama Japan. El Col1ins.P.. Akiyama T.. Shiraishi, I. and Nagase, T. “Services Expected for the Fit Phase of Space Tourism,” the 19th Inteanational Symposium on Space Xzhnology and Science, ISTS 94-g-25p, Yokohama Japan. Akiyama. T., “The Pleasure of Spaceflight.” the Journal of Space ‘lkchnology and Science, Vo1.9, No.1. 1993. The Sezond Meeting of the Medical Research Subcommittee of JRS Space Tourism at the Research Institute of Environmental Medicine of Nagoya University. January 20. 1994. PI Mitarai. G.. “Space Tourism and Space Medicine,” the Journal of Space Technology and Science, Vol.9, No.1, 1993. [lO]Hunter, Maxwell, W., ‘The SSX - A True
Spaceship,” Joumal of Practical Applications in Space, Vol.. No.1, Fall 1989 Issue. [ll]Koelle. Diih. E.. “Survey and Comparison of Winged Launch Vehicle Options,” the 19th Intanatid Symposium on Space Technology ad Science. ISTS 94-g-l Iv, Yokohama Japan. [12]Gaubatz. W-A., “Designing for Routine Space Access,” the Journal of Practical Applications in Space. Vol. N, Issue No.4. Summer 1993. [13]Grayson 0. and DiStefano E., “Propellant Acquisition for Single Stage Rocket Technology,” AIAA-93-2283.29th Joint Propulsion Conference and Exhibit, June 1993. [14]Koelle. D.E.. “BETA. a Single-stage Re-usable Ballistic Space Shuttle Concept.” Prccecdings of the Twenty-first Intcmadonal Astronautical Congress, North-Holland. 197 1. [lSlHudson, G.C., “Phoenix: a Commercial. Reusable Single-stage Launch Vehicle.” Pacilic American Launch Systems Inc., 1985. 1161Weeger, R.K., “Engine/Vehicle Integration far Vertical Takeoff and Landing Single Stage to Orbit Vehicle.” IAF-92-065 1.43rd Congress of the International Astronautical Federation, August 1992. [17jAkiyama,T., Personal Communication to the JRS Transportation Committee, February 1994, [18]Copper. J.A., “Single Stage Rocket Concept Selection and Design,” AIAA-92-1383, AIAA Space Programs and Technologies Conference, March 1992. [19]Kikukawa, H., Maruyama, T., “Current Status d R&D for Heat Resistant Structures for Spaceplane, the 19th International Symposium on Space Technology and Science, ISTS 94-b-17, Yokohama Japan. [201Torikai, T., “Space Tourism arxl Transportation,” the Journal of Space Technology and Science, Vol.9, No.1, 1993.
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CONSIDERATIONS ON VEHICLE DESIGN FOR SPACE TOURISM Kohki ISOZAKI,
Akira TANIUCHI and Koichi
$9.50+0.00
CRITERIA
YONEMOTO
KawasakiHeavy Induwies, Ltd. i Kawa.wki-cho,Kakand@hara-shi, Gijk 504, Jqxm
Hirosbige
KIKUKAWA and Tomoko MARUYAMA
Fuji Heavy Industries,Ltd. l-1-11 Yonan, Utsnnonajwshi, Tochigi 320, Jqm
The trauspoxtation research committee of JRS (Japaaese Rocket Society) has begun conceptual design of vertical takeoff and landing fully reusable SST0 (Single Stage to Orbit) rocke$ type vehicle as a smndard vehiile modelfffsplzemurian.Tlledesigncriteriaofthevehicle have paid most attention to the reqt&.ments of service to meet space tour amusement. Thestandard vehicle, which has 22m body length and weighs about 550 tons at takeoff, can provide attractive tours of 24 hours maximum for 50 passengers into the low earth orbit with a variety of w flight pleasures such as experience of weightlessness and earth sightseeing. Within the reach of our near future rocket technology, the design utilizes MMC, CF/Epy and Tiiw advanad matuials. The twelve LOX/LX2 engines consist of two nozzle types, which can be throttled ad gimbaied during the whole mission time, perform v&Cal launch and tail-fust reentry to final landing associated with raodynamic control of body flaps within tolemble accelaation acting on passengers.
The purpose of the project is to give a new, ad commercialization of space strong motivation for transportation by the space tourism studying from very different point of disciplines, such as space medicine, enterprise or tour service, which have never been learned in the history of Japanese rocket development (Elgure 1). Thus the aerospace industries are expected to make considerable efforts for producing low-cost spece transportation vehicles by the ground of seeking such a new space transportation market, namely space tour rref.21. [Jqnl”~eQ Rode1 society> Rsseorch
Grovps
Figure 1. Research Disciplines of Space Tourism The research project of ‘Space Tourism‘ has been adDpad for future sp&zeactivities by the Japanese Rocket Society (JRS). The study guideline was fast clarified by JRS Committee for Academic Activities at the annual meeting in April 1993. where experts in various disciplines wae invited to hold a pane1 discussion on this new theme [re.f.l]. The panelists were Mitarai Professor of Aerospace Physiology at Nagoya University, Kyotani Chairman of TFCNOVA. ‘R&ai Managing Director of JADC! (Japan aircraft Development Cotporation) and journalist Akiyama who experiured his space flight by Russian Soyu&Iir as the first Japanese astronaut. Copyright@ 1994 by the Intemational Astronautical F?edemtion.All right med.
223
The transportation research committee, the members of which are assigned from major Japanese BQOSPBC~ industries, airlines and corporate members of JRS interested in this project, was organized as a workiig group studying the definition of design criteria, the puspective of operational flexibility and finally a sladard vehicle model in cooperation with the other three rcscBcch activities such as medicine, enterp~se and tour service of spaze tourism (Figure 2). This paper summarizes the first year status repOn Of the research committee activity [ref.31including the issues of propulsion system design tref.41. liquid hydrogen technology rref.51.
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values of which may be completely different from current popular a&aft tour or virtual reality amusement
JRS Transportation Research Committee _%?f%ors. -Kawasaki
HeavyIndustries, Od.
21.
--
Mimbishi HeavyM&es.
-
Fuji HeavyIndustries,Ltd.
-
lshikawajimaH&ma HeavyIndutries, bd. Motor Co., Ud.
-Nissan -
AliNippon AirwaysCo., .bd.
-
Teisan KK.
Figure 2. Corporate Members of JRS Transportation Research Committee
.
.
2. Cuukbne
of %.w
Soace
Tour Model
Ltd,
The transportation research committee discussed two space tour models including amusement and available cabin services as shown in Table I. The flight plan is scheduled to take&f from a spaceport in northern Japan. The t~o-circles orbital flight course will provide a variety of m time, which offas Pacific Ocean / South Am&a sightseeing course at daytime depamue. Aliica / Asia sightseeing course at night depararre and so on.
Tourism
Table 1. Tour Course and Expected Service
Based on the guideline fist proposed by JRS, a unique study on the services expected for the first phase of space tourism [ref.@ will help u) clarify the features of space travel concerning tour course and cabin services similar to those of commercial airlines.
U-Hours Orbital Flight
BDqtime Depamtre * Pacific Ocean
. soulh America
Sight-
n Night
seeing
The key consideration on the vehicle design criteria for space tourism is how the ‘Spaceline’ vehicle can realize varieties of space flight joy that are most expected tour participants. A suggestion given by Akiyama was that the commercial values of the space tour = fust the experience of weightlessness and second the earth sightseeing [ref. 71.
~o-circlcs Orbital Flight
Departure . Africa . Asia
_ Amusement
Longitude
Depends on Departure Tie)
Weightlessness Astronautical Observation Telecommonication
_ Cabin Service
Recognition Souvenir
_
[ deg
Brcakfut/Lunch /Dinner
Soft Drink
Others The vehicle design pvided for space tourism should take these kinds of service scenario into consideration, the
mAI Amund the Eartl (DayandNight Scene
]
Figure 3. Two-&&s
Orbital Flight Tour
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As can be seen by the ground track of Figure 3. the vehicle leaving Japan flies first over the Pacific Ocean to Argentina. Then it passes the center of Africa to cross India and China. The passengers can enjoy spectacular scenery of their home town from the altitude of 200km. At the second mund of orbital flight the vehicle flies over Hawaii. The vehicle &orbits at At&a to pass the reentry inlerface at an altitude of KMkmin the middleof China. 2.2 Vehicle Design
Factors
Discussing the space tour in the low earth orbit as above is helpful to provide good engineering imaginations to designers. The design factors of vehicle for space tourism are C as follows.
cubin ArranPemetU The most important but difficult subjects that the vehicle designer should take into consideration ae passenger windows for sightseeing and a room for experiencing weightlessness. Discriminating the real sightseeing from the computer play like virtual reality, the vehicle should provide the windows as many as possible (Figure 4). The medical research group suggested that some passengers would have to observe the earth sitting on their own seats by physical reason, space sickness [ref.8].
Altitude Akiyama’s report accotding to his exciting joy flown by the Russian space station Mir is meaningful. He commented that the higher orbit altitude does not necemarhy deserve to raise the value of space tour service [ref.7]. Thus the ahitude of 200km is assumed appropriate for the sightseeing orbit. Orbit Inclination Since the earth’s equatorial region is sparsely populated, as we know, most of the spaceports will be constructed at the northern hemisphere at relative high latitude rref.61. The locality of population is one of the major factors of business to determine the orbit inclination. Although launching from higher latitude has disadvantage in payload injection capability, it provides passengers attractive earth sceneries that cover the equatorial zone to the both poles.
Figure 4. Window Arrangement A microgravity amusement space will be necessary for passengers who will enjoy floating in microgravity condition preventing accidents from kicking other’s head
Tour Time The maximum flight time of 24 hours can be preliminary determined by the tour requirement to cover whole sight of the earth in the daytime. This maximum time is also a trajectory restriction returning to the home spaceports located at relatively high latitude. On the other hand, the minimum flight time that the transportation reaearch committee has planed is two circles around the earth. This minimum time came tiom the necessity of preparine deorbit chance. Since one third of each orbital time is in d&mess. the m time has significant meaning for selection of sightseeing course.
Figure 5. Microgravity Amusing The idea of facilities currently used by airlines such as galley, toilet, miscellaneous utilities. television screen for monitoring scenery and sightseeing commentary system will be helpful in designing cabin arrangement
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. .
al Restrrctrons The passengers suffer most critical acceleration at ascent phase to orbit. As is already recommended, the vehicle shouldn’t exceed the acceleration limit of 4 G from head to foot and 2 G in the reverse way tref.91. Taking account of the medical endurable load and its direction in mind, more relaxed accelemtion level should be con&&d to prevent passengers from blackout.
Another medical problem is space sickness. Many passengersare thought to suffer severe space sickness as soon as they areput in weightlessness. Although there are medical treatments such as oral or transdermal medicine. mechanical methods like LBNP (Lower Body Negative Pressure) or Seat head rest (Figure 6) would be also effective as the first aid [ref.8]. Head Rest /
\
=P
, which have smaller cross range capability than the nose-first ones, don’t require high attitude rotation maneuver. Giving the design feasibility for passenger priority to accommodation, the uansportation research committee adopts the tail-first reentry style. 3.2 Performance
Reauirement
The total velocity increment required by orbital conditions is summarized in Table 2. The orbitai altitude is 2OOkm with the inclination of 45 degrees. Velocity increment of 300m/s is assumed for the vertical landing maneuver and 290m/s is added to the total velocity as the performance margin. The resultant total velocity of 9.93km/s is relative large compared with that required for BETAbef.141 or Phoenix[ref.lS]. Table 2. Total Velocity Increment
Figure6. PassengerSeat Design Mission 3.
Phase Velocity Increment
Standard Vehicle Orbital Velocity Gravity Loss Drag Loss Maneuver Loss Deorbit Impulsive Velocity Descent Landing Maneuver ____________________--Design Margin Aseem
Although winged launch systems are attractive ad mentally me in expectationfor futurecommercial space transportationstyle on the analogy of current &l&s, vuticai takeoff and landingfuily reusableSST0 (Single Stage to Orbit) mcke4 type vehiiles are generally thought feaaiblc within near-term technology [nf.lOl [rcf.ll]. Thus the muqwution nseareh cunmittee adqwdthis
vutical mkeoffandlandingSSTO
thestmtddvehiclemodelforthepwentsp8cetourism study.
Total Velocity
&m/s1 7.701 0.900 0.600 0.070 0.070 0.300 _____0.290 9.930
rocket=
2.3 Confiwration
Design
Since he design criteria of the vehicle for space tourism dependmuch on the user’srequestof sightseeing and weightlessness experience. the cabin arrangemauhas significant influence on the vehicle’s configuration. The vehicle design results in the body length of 22m with the bottom diameter of 18m as shown in Figure 7.
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22m
k-----4
Figure 7. SST0 Vehicle for Space Tourism
lgm------/
A two-floors cabin equipped with many sightseeing windows and a microgravity amusement space makes the top of the vehicle in upright position. The propellant tank is a semi-integral structure using advanced material to reduce weight, which has common bulk head between LH2 and LOX propellant tanks. livelve e@nes. which consist of four booster ad eight sustainer engines with conventional bell nozzle. ae mounted in a cituular position around the lower tank structure. Noxakexpansion ratios are 15 for the bwster engines and 40 to 80 (two positions) for the sustah~~ eWmes.nspectivdy. The vehicle has four l&age type landing gears, which can be -ted in the body. The energy absorption concept is conventional oleo pneumatic system. Ib prevent topple &rum of the vehicle in case of cue kmding gear failure, the length of oleo stroke in the opposite position will be s&x&ted. Delta Clipper has pvsed an idea to extuxl the noxxle of a sustainer engine instead of tbehtikdlandinggear[ref.l6].
3.4 Cabin Figure 8 shows the comparison of two typos of cabin arrangement. One lines up passenger seats straight like current airlines and the other adupts the seats lined up in circle to provide better view through windows. Akiyama arrsnganentincheleasthebu&ine -theseat [ref.l’ll. An image of cabin amtngaWtwith~gbtlinedup seatsisreahzedbyascalemodelsbmvninFii9.Tbe microgravity amusement space surrounded by a tmnqamntwallwithanentrancedoor,whi&~totbe upperdeck,locatesattbenarpositionoftheubinflooc l!vo waste management compartments ain seen in the [email protected] flightofstairstotheupperdeckisalsore&ed. lbereamtwopilotseatsatthefmntsideoftheupper deck, from which enough wide view is available chrting reentry and landing operation (Figme 10). Aa airlo& seen in Figure 1 I is also equipped on the upper deck m emergency exit.
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Entrance
Cockpit
Door
Alflock
Amusement Space
UPPER DECK
In-line Seat Amneement
a: cockpit
PIlot Seat
Galley
Alh& /
Micmgravity Amusement space
8. Cabin Design
Stewardess
Seal
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Mass Charactcristic$
The mass characteristics of the present vehicle estimated by the first design work is shown in Table 3. Table 3. Mass Characteristics Subsystems 1. E%ucmrc ~Engincskirt
9Engine lluust smrc * Maincabiisaucture - &ckpitK.abii Smtcrure
2.
3.
4. 5. 6. 7. 8.
* ‘Ihamal Pm&don System Heat Shield (Including Sustainer EIG Cover) * Landing Gear System * Airlock System Prupulsion System * Sustainer Engine * Booster Engine * LH2-Tank * LOX-Tank . AuxiliaryTanks (LH2,LoX) . FkwuizationSystem (AHe and CHe) . Reaction Control System . Feed System . Residual Prolxllant (Excluded in Subtotal) Actuatca System . Auxiliary Power Unit . h4isccUancous Actuators and Pumps Environmental Conuol System Power Supply System Navigation, Guidance and Control System Communication and Data Acquisition System Misccllanwus Equipment * Passenger Seats * Crw Seats * Galley * Toilets
* Misauancous
Subtotal kg] 15,959 3,860 2,063 2,009 2,436 3,944 1.375 272 29.5 17 9,456 4,184 8,888 4,089 459 1.950 400 100 2,475 820 220 600 1.100 580 305 431 1,542 1.127 62 122 130 100
Empty Mass 9. Passengers and Crew Weight * Fiity Passengers 3,750 * crw (FourPersons) 300 * Luggage 270 10. Propellants (Including Residual Propellant) 1Fuel (LHZ) 70.703 *oxid&r(L0X) 424.2 15 Total Lift-off Mass (Including Design Margin of 597kg)
SO.254 [kg1 4,320
494,918
550,089 [kg)
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4. S-cm 4.1
Acrodvnamic
Since the slush hydrogen will have unknown system complexity in the propellant feeding. the Norman cryogenic hydrogen became the baseline. The major disadvantages using plug nozzle are considered the lack of roll control force and the inflexibility of attitude control for engine-out contingency. Thus the conventional bell nozzle type with throttle seems feasible for the the st&ard vehicle.
Desim
Des&n
The standard vehicle employs four body flaps aI the engine skirt section, the two windward flaps of which activate as’ angle-of-attack trim control and the two leeward flaps are utilized for lateral stability. In order to evaluate the lift-t&rag characteristics and the trim capability, aerodynamic analysis is performed using the calculation model shown in the left hand side of Figure 12.
The transportation research committee has spend a long discus&u to configure booster and sustainez engine numbers, the cycleof which is finally &&Mto employ ‘augmented expander’. There are two reasons to decide the lotal engine number. The first reason is to keep Uuuat-toweight ratio at liftoff between 1.3 to 1.5 using 80% thrust level of LE-7 engine enhancing reliability and life cycle. The another reason is so-called intact abort criteria to maintain the thrust-to-weight ratio more than one in case of two engine failure [ref.lO].
As a result shown in Figure 12, the vehicle can attain the lift-to-drag ratio of 0.3 and 0.4 at the an&+f-attack 20 and 35 degrees respectively at hypersonic flight. Since the lifting body characteristics is about one third of the current performance of Space Shuttle and lower than the predicted Delta Clipper capability [ref.18], this type of tail-fust reentry vehicle may not have enough cross range modulation capability to return from high inclination orbits to contingency bases.
9.2 PrmUon
The expansion ratio for the booster engine was selected 15 to drive maximum thrust at liftoff. Due to flow separation limit of engine nozzle at sea level, on the other hand, the expansion ratio of 40 for the nzb~ted nozzle position of sustainer engine is feasible. The sustainer engine realizes Ihe expansion ratio of 80 at the full extended position within available space and weight. The propulsion group did enthusiastic study to find out optimal combination of the booster and the sustainer engines [ref.4]. The summary is shown in Table 4.
Design
The study on propulsion system has first started with the discussion whahex the standad vehicle adopts normal cryogenic hydmgen propellant rather than slush hydrogen, and whether it should employ conventional bell notie instead of plug nozzle type engines.
.s
B 4
..-
..___
g
‘-
J
9”
-a”
40 -30 -20 -10 Angle-of-Attack (deg)
0
10
0.04 E .s g 0.02 s z
0
E z0 -0.02 P E -0.04 0 ii 6.06 -60
Fiaurc 12. Lift-to-Draa..mand Trim _~ ___.__.____ ~~ Char~teristicn
-50
40 -30 -20 -10 Angle-of-Attack (deg)
0
10
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Tkble 4. Engine Configuration lI&otT Engine Characteristics
Note) + fiied expansion ratio of 40 Asthenumbtrofsusmineratginehuzases,theorbit injection cqability gets significant gain including weight penalty spend for the extension nozzle equipment. This payload gain is due to the long duration of sustainer engine operatitm with higher specific impulse. Since the vehicle’s base illt4 is restricted, there is a limitation of sustainer engine mrmber. The sUmlard vehicle adopts a combination of four booster engines and eight sustainer engines as the baseline (Figure 13). During terminal flight and laadiag phase the two booster engines ae Wivated while the other two rue put in its idle mods preparing for engine faihue. Suruiner mine
Acceleration of rotation required for the rwcticn control system is 1 deg/s*. which results in about 7OOON class thruster using GOX/GHZ propellant. The total impulse requirement of 4.2mNs was e&rated to satisfy rotation duty for the earth sightseeing. lhnk baffle, propellant feed system, pressurixation and vent system at alsoiXWKqWUystudied.
One of the critical design issues and requirements far realizing the SST0 vehicle for space tourism is the structural design. In order to minimize inert mass, various material candidates and structure system have been surveyed as shown in Figure 14. which wereoriginally studied for application of futute ae!rqXe plane [ref.191.
)66@3y
BOOSILX Engine
0
I
O
0
It is imbcated that the present technology hasalmost luIchedtothetargetmasslevelfortheprimarysuuctme, but more development effort is needed to realize the cryogenic tank structure.
Structure materials applied to the present staWd vehicle am seleaed among the liiht weight combinatiats of conventialal TPS (TbenMl FVoteuion System) and the advamM structure system mentioned above llccadins to the thermal requirement. The result of the Mine structure is shown in Tbble 5 and Figure 14. 18m _____I Figure 13. Engine Anangemalt at Bottom of vehiile
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Unit Mass
Unit Mass
ko/m’
0
kC/m 2
4 g s 1 _.__._-.._ ........_.. 500
lti
1SiX
0.
2ooo
Temerature (Kl
Figure 14. AdvancedSe~cture Survey ‘IWe 5. Tkmal Condition of Baseline Structure Smkxu-dl ComponcnrlArea
Surface Temperature
Heat Shield
1400-1100
MMC + SA/HC(TPS)
Engine Skin
1100-800
MMC + WHCIJ-PS) Al Alloy + Tiiw(-rPS)
Up to 800
CdbiiMOS.5
4.4 Nor-.
Material& Concept
(K)
CF/Epoxy/HC+
LH2-T?lIlk
ml
LOX-Tank
Up to 800
h0.K
-Metal Mm-ix
SA/HC
-Supa Alloy I Honeycomb
Composiu
Tiiw
Avionics of navigation, guidance and control system for the present SST0 vehicle has no aitical design issue. Attention will be paid for the onboard software design for both the ascent and reentry phase that calculates not only the optimal trajectory guidance and control but also manage various fail- modes as safe as possible [ref.u)].
Tiiw(-l-PS)
+ Insulation CF/Epoxy/HC+
-Titanium
CF/Epoxy/HC
Guidance and Control Svstem
+ Intcmal Insuladon
I Multi-
-Carbon Fitul
4.5
Insulation
Wall Epoxy /Honeycomb
ECLSS
ECLSS (Environmental Control and Life Support System) has significant meaning to provide many passengers comfortable space flight. A redundant system concept is presented in Figure 13.
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Figure 15. Structural Concept of the Standard Vehicle C__________________________._________-_~-------_______~
THC:Tempe-mdHumidirycOnd ARs:hmaphcrsRevipliudmspw .CO2Rmal .TKS C,mmhmnt Conml
Figure 16. ECLSS for the Standard Vehicle
4.6 Powcr Su~dv
.Cs:f-&~=-&w~Y ACM :Atmqhm
Comqc4ih
?.b”imtil
FDS : Fus Dction amiSuppmsion
1000.0T
Svstcm
The present SST0 vehicle will be equipped with two
kinds of power supply systems, which are an electrical power system and a hydraulic power system. The electrical power system is comprised of Fuel Cell Systems that use cryogenically stored oxygen ad hydrogen reactants, Reactant Storage/Distribution Systems and Electrical Control Units. The electrical power system produces all the electrical power requked by ECLSS, cabin services and other onboard avionics equipment for the entire flight duration.
z
5:o.o
Altitude [km]
500.0 z %
$0.0 0.0
Altitude [km]
Figure 17. Engine Char~teristics
I
30.0
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z
--
E -i+ F
---_
0.0
3
2
[s]
a
I
0.0
Tfne
0.0
Time [s]
350.0
‘Time [s]‘
, 350.0
350.0
Relative Velocity [km/s]
700.0 ?
0.0 8.0
1
i
/ 0.0 00 250.0
Relative Velocity [km/s]
T
00
Time [s]
1000.0
T
,-------r
.i 0.0
lime
[s]
0 Figure 18. Ascent Trajectoav
350.0
1000.0
f
I 00
0.0
:
:
;L
Time [s]
Time (s]
Figure 19. Reentry Trajectory
1000.0
350.0
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it possible to apply very light-weight material for protecting the thermal load on the base surface. 200
One of the critical performances for this kind of SST0 vehicle is the cross range capability. m calculation predicts that the cross range achieved by the aemdynamicbankmodulationwiU healitUelatgerthan 200km (Figure 20). Although the vehicle can not p&rut high maneuver in order to land any various m this will be huge enough to assure emegcllcy landing to contingency spaceports.
-S 5 100 8 4S
0
$ -100 0e -200 300
t -
’
’
’
’
2000
1000
’ k ’ ’ ’ *’ 4000 3000 Down Range [km] ’
*
’
’ ’ 5000
Figure 20. Range Modulation Capability For high intensity, but rather shorter duration power requirements that is called for by rocket engine gimbal actuation systems or by the body flap actuators, the hydraulic power system provides the necessary hydraulic power which is generated either by Auxiliary Power Units or by the rocket engine driven hydraulic pumps. In the case when the APUs are to be used, those will be opemted with the hydrogen/oxygen fuel to ensure the environmental friendliness. The APUs will be opemted only during the ascent and descent phase of the flight Operation.
5.
Traiectorv
Analvsis
As shown by the engine characteristics (Figure 17), the sustainer engines extend their nozzles to the seccnd position when the specific impulse of expansion ratio 80 exceeds that of 40 near the altitude of IOkm for ascent phase. A typical trajectory is simulated as shown in ‘Figure 18. The four booster engines begin throttling when the vehicle reaches the acceleration level of 3 G at about 100 seconds after liftoff. The vehicle achieves the altitude of 200km in 6 minutes. trv Tru
The guidancesimulating the reentry trajectory uses a conventional algorithm of drag acceleration control widely applied to winged vehicles. Since the acceleration acting on passeargus is limited 3 G. the nominal reentry load factor considering safety margin is designed 2.5. A typical reentry trajectory is shown in Figure 19. The low heat input due to the relative small ballistic coeffiiient and a large radius of the base curvature makes
. 6. Coneludlnn Studies on SST0 vehicle for commercialization of space tourism is about to start. The mn research committee was organized to define a sgldsld vehicle that is expected to be a real model for the space tour flight. It held nine meetings in total discussing on the vehicle design for the first yea. The committee has finally build a l/20 scale mockup shown in Figure 21. There have been many issues newly found md discussed at every meeting. Wb have rea@zed that are of the most critical technology challenges is mataiak Although the standard vehicle adopts various adnrred light-weight materials as much as possible. we had to accept 15% mass reduction horn the estimation hased on the cunent technology. Wegraduallyunderstandthatwehavetodumpourold fashioned way of thinking. Wb will conclude the present paper with the question “Can the enthuaiaam for Space Tourism save Japanese ‘stereotyped’aem+g~~ induauies ?” Probably it will be answemd by “Yes” in the nex future.
The authors are grateful to the members of the JRS Transportation Research Committee for space tourism, and especially Prof. M. Nagatomo and Mt. Y. Naruo of the Institute of Space and Astronautical Science, Mr. T. Ruikai of Japan Aircraft Development Agency and IL: P. Collins of National Aerospace Laboratory for their kind advises and useful discussions.
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Figure 21. Mockup of SST0 Vehicle for Space Tourism
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Referentas
(11 Nagatomo, M., “On JRS Space Tourism Study Program,” the Journal of Space Technology and Science, Vol.9. No.1. 1993. PI Collins, P., “‘IbwatdsCommercial Space Travel,” the Journal of Space Rchnology and Science. Vo1.9, No.1, 1993. r31 Isozaki, K., lhniuchi. A., Yonemoto, K., Kikukawa, H., Matuyama T.. Asai, T. and Murakami. K., “Vehicle Design for Space Tourism,” the 19th International Symposium on Space Technology ad Science, ISTS 94-g-22p, Yokohama Japan. [41 Mori, K.. Suzuki, A.. Iihara, S. and Nakai, S.. “Requirements on Propulsion System Design and Operation for Space Tourist Car&r Vehicles,” the 19th International Symposium on Space Technology sod Science, ISTS 94-g-23p. Yokohama Japan. El Hanad&T., Nagatomo, M. sod Naruo, Y., “Liquid Hydrogen Indusuy : A Key for Space ‘Iburism,” the 19th Intanational Symposium on Space Technology and Science, ISTS 94-g-24p. Yokohama Japan. El Col1ins.P.. Akiyama T.. Shiraishi, I. and Nagase, T. “Services Expected for the Fit Phase of Space Tourism,” the 19th Inteanational Symposium on Space Xzhnology and Science, ISTS 94-g-25p, Yokohama Japan. Akiyama. T., “The Pleasure of Spaceflight.” the Journal of Space ‘lkchnology and Science, Vo1.9, No.1. 1993. The Sezond Meeting of the Medical Research Subcommittee of JRS Space Tourism at the Research Institute of Environmental Medicine of Nagoya University. January 20. 1994. PI Mitarai. G.. “Space Tourism and Space Medicine,” the Journal of Space Technology and Science, Vol.9, No.1, 1993. [lO]Hunter, Maxwell, W., ‘The SSX - A True
Spaceship,” Joumal of Practical Applications in Space, Vol.. No.1, Fall 1989 Issue. [ll]Koelle. Diih. E.. “Survey and Comparison of Winged Launch Vehicle Options,” the 19th Intanatid Symposium on Space Technology ad Science. ISTS 94-g-l Iv, Yokohama Japan. [12]Gaubatz. W-A., “Designing for Routine Space Access,” the Journal of Practical Applications in Space. Vol. N, Issue No.4. Summer 1993. [13]Grayson 0. and DiStefano E., “Propellant Acquisition for Single Stage Rocket Technology,” AIAA-93-2283.29th Joint Propulsion Conference and Exhibit, June 1993. [14]Koelle. D.E.. “BETA. a Single-stage Re-usable Ballistic Space Shuttle Concept.” Prccecdings of the Twenty-first Intcmadonal Astronautical Congress, North-Holland. 197 1. [lSlHudson, G.C., “Phoenix: a Commercial. Reusable Single-stage Launch Vehicle.” Pacilic American Launch Systems Inc., 1985. 1161Weeger, R.K., “Engine/Vehicle Integration far Vertical Takeoff and Landing Single Stage to Orbit Vehicle.” IAF-92-065 1.43rd Congress of the International Astronautical Federation, August 1992. [17jAkiyama,T., Personal Communication to the JRS Transportation Committee, February 1994, [18]Copper. J.A., “Single Stage Rocket Concept Selection and Design,” AIAA-92-1383, AIAA Space Programs and Technologies Conference, March 1992. [19]Kikukawa, H., Maruyama, T., “Current Status d R&D for Heat Resistant Structures for Spaceplane, the 19th International Symposium on Space Technology and Science, ISTS 94-b-17, Yokohama Japan. [201Torikai, T., “Space Tourism arxl Transportation,” the Journal of Space Technology and Science, Vol.9, No.1, 1993.