In 1993, the Japanese Rocket Society ( JRS) adopted space tourism as a research project for future space activities of mankind (1). The purpose of the research is to give a new and strong motivation for commercialization of space transportation by studying space tourism from various viewpoints of different disciplines, such as space medicine, business or passenger service. Such a study has never been made in the history of Japanese rocket development.

As discussed (2), aerospace companies are expected to make considerable efforts to produce lowcost space transportation vehicles on the grounds of seeking a new space transportation market, namely space tours. Since most people have no doubt about the fascination of flight to orbit, only the 'how' and 'when' of its realization are open to question. Spreading the market to establish regular 'Spacelines' from ground to orbit will result in a miraculous cost reduction in space vehicle production and flight operations.

The transportation research committee of JRS was established to study real space vehicles which could be used for tourism. The committee members consisted mainly of the corporate members interested in this subject. As a result of regular workshops of the committee, this paper summarizes the first year status report of the standard design criteria for the vehicle and its operation. The propulsion system design (3) and related study on liquid hydrogen technology (4) are included in this issue.

Besides a guideline proposed by the JRS Committee for Academic Activities, the first phase space tourism study (5) was used to clarify the features of space travel concerning tour courses and cabin services similar to those provided by commercial airlines. In order to define a vehicle concept for space tourism, a vertical takeoff and landing fully reusable SSTO (Single Stage to Orbit) rocket vehicle was chosen without a detail tradeoff with aerospace planes. This type of vehicle has been studied by several authors (refs. 6,7,8 & 9). For the size of the vehicle, this study assumed that the liftoff mass should be equal to the take-off mass of a typical wide-body airplane.

According to a comment on the pleasure of spaceflight by the first Japanese astronaut, Akiyama, the basic commercial values of a space tour are the experience of weightlessness and Earth sightseeing (10), and further considerations on space tourism services have been made for the JRS Space Tourism study (1). Based on these comments and considerations, Table 1 summarizes the types of services which space tourists can expect in an early space tour. The space vehicle is expected to provide passengers with these unique experiences which differ completely from the current virtual reality amusement.

Table 1 Space Tour Service Model

In Table 1, the first standard flights are assumed to be two of orbit flights, and the second ones would last for twenty-four hours. As can be seen by the ground track of Figure 1, a 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. On the second orbit the vehicle flies over Hawaii. The vehicle then deorbits over Africa to pass the reentry interface at an altitude of 100km in the middle of China.

Such a scenario implies technical requirements of vehicles to be used for space tourism in the near future. The following are typical design factors to be considered.

According to Akiyama's comments, flying at higher altitude does not necessarily provide better sightseeing, based on his flight in the space station Mir. Thus, the altitude of 200km is assumed to be our standard orbit for sightseeing. No orbit transfer capability will be provided.

Usually, a space vehicle can carry a heavier payload when it takes off eastward from lower latitude. In this respect, the equatorial zone is the best site for spaceports. However, higher inclination orbits have an advantage in providing passengers with a wider range of more attractive views of Earth from the equator to the high latitudes, at the sacrifice of payload injection capability depending on spaceport locations. A space vehicle for space tourism should have enough launch capability to satisfy this demand and to match the population distribution of prospective tourists.

The maximum flight time of 24 hours can be preliminarily determined by the tour requirement to provide views of the whole Earth in daytime. This maximum time is also set by trajectory restrictions for return to takeoff sites located at relatively high latitude.

On the other hand, the minimum flight time that the transportation research committee has planned is two orbits. This minimum time came from the necessity of preparing for deorbit. Since one third of each orbit time is in darkness, the departure time is significant for selection of sightseeing course.

Two important but difficult requirements that the vehicle designer should take into consideration are windows for sightseeing and a room for experiencing weightlessness. In order to distinguish real space sightseeing from computer play like virtual reality, the vehicle provides as many windows as possible. The medical research group suggested that some passengers may have to observe the earth sitting in their own seats for physical reason such as space sickness (11). A microgravity amusement space will be necessary for passengers to enjoy floating in weightlessness safely.

The ideas of facilities currently used by airlines such as galley, toilet, miscellaneous utilities, television screen for monitoring scenery and sightseeing commentary system will also be helpful in designing the cabin arrangement.

Passengers will experience the greatest acceleration during ascent to orbit and atmospheric reentry. Although it was requested that the acceleration shouldn't exceed 4 G from head to foot and 2 G in the reverse way, a stricter acceleration level of 3G from head to foot and 0 G in the reverse direction was used for this design study (12).

Another medical problem is space sickness. Although there is no standard to design the vehicle against space sickness, provisions for future medical treatments and first aid will also be used in designing the cabin.

The configurations of SSTO rockets are classified by atmospheric reentry style into two types; nose-first reentry configurations and tail-first reentry configurations. Vehicles that perform the nose-first reentry have a relatively slender body like the Delta Clipper (9), which reduces the aerodynamic drag to achieve better orbit injection capability and attains relatively large cross range capability during the reentry phase. However, nose-first entry vehicles require a maneuver to rotate 180 degrees in the atmosphere before they land vertically. This maneuver brings difficult problems to vehicle design in terms of acceleration control, that are critical for passenger accommodation and for surface control of liquids in propellant tanks (13).

On the other hand, tail-first reentry vehicles like the BETA (7) and the Phoenix (8) can avoid these problems, although the cross range capability for this type is considered to be smaller than nose-first reentry vehicles. From this view point, a tail-first type of vehicle has been adopted for

The orbital condition and total velocity requirement are summarized in Table 2. The orbital altitude is 200km with inclination of 45 degrees. A 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 relatively large compared with that for the BETA or the Phoenix.

The design of vehicles for space tourism depends heavily on the users' request of sightseeing and weightlessness experience. Thus the cabin arrangement has much influence on the configuration of the vehicle. The design results in a body length of 22m with a bottom diameter of 1 8m as shown in Figure 2.

A two floor cabin 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 a common bulkhead between LH2 and LOX propellant tanks.

Twelve engines, four booster engines and eight sustainer engines, with conventional bell nozzles are mounted in a circular position around the lower tank structure. Nozzle expansion ratios are 15 for booster engines and 40 and 80 (two positions) for sustainer engines, respectively.

The vehicle has four linkage type landing gears, which can be retracted in the body. The energy absorption concept is a conventional oleo pneumatic system. To prevent toppling of the vehicle in case of one landing gear failure, the length of the opposite oleo stroke is shortened.

Figure 3 shows the comparison of two cabin arrangements that were considered. One lines up passenger seats parallel like current airlines (above) and the other, recommended as better by astronaut Akiyama (14) adopts the seats arranged in circles to provide better views through the windows (bottom).

The mass characteristics of the present vehicle estimated in the first phase design work is shown in Table 3. It should be noted that the total lift-off mass exceeds the target mass of a widebody aircraft.

	Subsystem / ¥ Items	Subtotal
1.	Structure	15,959
	¥ Engine Skirt	3,860
	¥ Engine Thrust Structure	2,063
	¥ Main Cabin Structure	2,009
	¥ Cockpit/Cabin Structure	2,436
	¥ Thermal Protection System Heat Shield	3,944
	(including Sustainer E/G Cover)
	¥ Landing Gear System	1,375
	¥ Airlock System	272
2.	Propulsion System	29,517
	¥ Sustainer engine	9,456
	¥ Booster Engine	4,184
	¥ LH2-Tank	8,888
	¥ LOX-Tank	4,089
	¥ Auxiliary Tanks ( LH2 and LOX)	450
	¥ Pressurization System (AHe and CHe)	1,950
	¥ Reaction Control System	400
	¥ Feed System	100
	¥ Residual Propellant	2,475
	(excluded in subtotal)
3.	Actuator System	820
	¥ Auxiliary Power Unit	220
	¥ Miscellaneous Actuators and Pumps	600
4.	Environmental Control System	1,100
5.	Power Supply System	580
6.	Navigation, Guidance and Control System	305
7.	Communication and Data Acquisition System	431
8.	Miscellaneous Equipment	1,542
	¥ Passenger Seats	1,127
	¥ Crew Seats	62
	¥ Galley	122
	¥ Toilets	130
	¥ Miscellaneous	100
Empty Vehicle Mass		50,254
9.	Passengers and Crew Weight	4,320
	¥ Fifty Passengers	3,750
	¥ Crew (four persons)	300
	¥ Luggage	270
10.	Propellants	494,918
	(including Residual Propellant)
	¥ Fuel ( LH2)	70,703
	¥ Oxidizer ( LOX)	424,215
Total Lift-off Mass (including design margin of 597 kg)		550,089

Aerodynamic design of this vehicle is mainly concerned with aerodynamic stability and cross range control during the atmospheric reentry phase. The present vehicle employs four body flaps at the engine skirt section, the two windward flaps of which act as angle-of-attack trim control, and the two leeward flaps are utilized for lateral stability. A calculation model shown on the right hand side of Figure 4 was used for the aerodynamic analysis of the vehicle.

As a result shown by Figure 4, the vehicle can attain lift-to-drag ratios of 0.3 and 0.4 at angle-of-attacks of 20 and 35 degrees respectively in hypersonic flight. Since the lifting body characteristics are about one third of the current performance of Space Shuttle and lower than the predicted Delta Clipper capability (15), it may not fully satisfy the cross range requirement to return from high inclination orbits on a contingency basis.

At the beginning of the vehicle design for this study, the requirements for the propulsion system have been analyzed. Consequently, a standard type of cryogenic propulsion system using liquid hydrogen and liquid oxygen as propellants rather than slush hydrogen has been selected due to their wellknown characteristic. For a similar reason as well as for flexibility in engine-out contingencies, the tradeoff study of engine nozzle types resulted in selection of bell nozzles rather than a plug nozzle. The number of booster and sustainer engines has been determined to fulfill the requirement of total thrust-to-weight ratio at liftoff between 1.3 to 1.5, and intact abort criteria to keep the thrust-to-weight ratio more than one in case of two engine failures, based on practical data using 80% thrust level of LE-7 engine enhancing reliability and life cycle. Although the orbit injection capability would be increased with more sustainer engines, a combination of four booster engines and eight sustainer engines has been adopted due to the limitation of the vehicle base area shown in Figure 5. For the terminal powered flight during landing, two booster engines will be activated while the other two are put in idle mode preparing for engine failure.

Angular acceleration required for the reaction control system is 1 deg/s2, which results in about 7000N class thruster using GOX/GH2 propellant. The total impulse requirement of 4.2mNs was estimated to satisfy the rotation duty needed for Earth sightseeing.

The most important requirement of structural design is to minimize the mass of propellant tanks containing low density liquid hydrogen propellant, and of the sophisticated reusable thermal protection, which are critical for the performance of SSTO vehicles.

This design study is based on another study (16), which compared various candidate materials and structural systems originally developed for future aerospace planes, in terms of a performance parameter of "unit mass" defined by structure mass per structure surface area. The result is summarized in Figures 6 and 7 for the primary structure and integral tank structure, respectively. The unit mass targets specified in both Figures are based on design requirements of various vehicles. Figure 6 shows that present technology can realize the targets for the primary structure. On the other hand, present technology cannot satisfy the requirement for the integral tank structure as shown in Figure 7. It will be necessary to design the tank geometry to minimize the surface area of the constant volume tank.

Considering not only the technical but also the economic aspects of these structures, we have designed structural systems for the present vehicle with light weight combinations of conventional TPS (Thermal Protection System) and advanced structural system to meet the thermal conditions of various parts of the vehicle as shown in Figure 8.

Avionics for the navigation, guidance and control system for the present SSTO vehicle has no critical design issue. Attention will be paid to the onboard software design for both the ascent and reentry phase to calculate not only the optimal trajectory guidance and control, but also to manage various failure modes safely (17).

The ECLSS (Environmental Control and Life Support System) has a significant role to provide many Passengers with a comfortable space flight. A redundant system concept is presented in Figure 9.

The present SSTO 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 and hydrogen reactants, Reactant Storage and Distribution Systems and Electrical Control Units. The electrical power system produces all the electrical power required by ECLSS, cabin services and other onboard avionics equipment for the entire flight duration. 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 operated with hydrogen/oxygen fuel to ensure environmental friendliness. The APUs will be operated only during the ascent and descent phase of the flight operation.

A typical ascent flight has been simulated and a time sequence of flight conditions is shown in Figure 10. The four booster engines begin throttling when the vehicle reaches the acceleration level of 3 G about 100 seconds after lift-off when the vehicle gain has gone only a small downrange distance and is within safe distance for intact recovery at the launch site. The vehicle achieves the altitude of 200km in 6 minutes. More detailed analysis will be required to determine emergency flight plans in case of failure of the propulsion system.

The guidance for the reentry trajectory simulation uses a conventional algorithm to control acceleration, which is widely applied to winged vehicles. Since the acceleration acting on passengers is limited to 3 G, the nominal reentry load factor including a safety margin is designed to be 2.5G. A typical reentry condition is shown in Figure 10. The low heat input due to the relatively small ballistic coefficient and the large radius of the base configuration makes it possible to apply very light-weight materials for thermal protection to the base surface.

One of the critical performance parameters for the present SSTO vehicle is the cross range capability. Our calculations predict that the cross range achieved by aerodynamic bank modulation will be a little more than 200km (Figure 11), which will limit orbit inclination and spaceport location, but will be large enough to assure emergency landing at medium latitude spaceports.

Figure 12. Upper part of scale model built for conceptual design study of an SSTO for space tourism.

Studies on SSTO vehicles for commercialization of space tourism are about to start. The JRS transportation research committee was organized to define a vehicle model which was expected to be used for related studies as a reference vehicle. It held nine meetings discussing the feasibility of the vehicle design and available technologies over the past year.

As a preliminary result, a concept of a passenger vehicle to carry fifty passengers for a two orbit flight in low earth orbit has been developed. A scale model of the vehicle has also been built (Figure 12). However, many technology challenges were found at each meeting. In particular in the field of materials, we had to assume a 15% mass reduction from the current estimate for the structure through using advanced materials, such as CF/Epoxy/HC or MMC. Reviewing the current design and examining details of the scale model design, it seems to be possible to improve the vehicle design by further efforts.

Thus, we gradually understood that we will have to dump our old-fashioned way of thinking. We will conclude this paper with the question "Can enthusiasm for Space Tourism revive the 'stereotyped' approach of Japanese aerospace industries ?" Probably it will be answered "Yes" in the near future.

The authors are grateful to the members of the JRS Transportation Research Committee for space tourism, especially to Prof. M. Nagatomo, Mr. Y. Naruo, Mr. T. Torikai and Dr. P. Collins for useful discussions.

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