Space Tourism is the only foreseeable market area where a large and growing demand can be expected in the future. This has been confirmed by several studies, ref.1.

The condition, however, is the availability of launch vehicles and an operations scheme that allows ticket prices of less than 50 000 $ at launch rates of several 100 flights per year. None of the present or planned vehicles ( such as Venture Star") can meet this requirement. A completely new approach is required, strictly according to the rules of Cost Engineering", i.e. all programmatic, vehicle design and operational features have to be optimized with the goal of minimum cost.

It may be helpful in this respect that a space tourism vehicle development (and operation) is a strictly commercial venture. It is not a task of governmental space organizations.

Otherwise, if an international consortium for the development of a dedicated space tourism vehicle could be established then the high-frequency operations would allow to offer traditional space transportation services for prices which are another order of magnitude lower than those expected from the first generation of reusable cargo vehicles at present launch rates.

This means space tourism as such is a new market but it would pave the way for many other space applications (such as space power stations, by example). Only this further cost reduction may also allow the eco-nomic establishment and operation of an orbital hotel facility.

Several options for space tourism can be defined:

1. Ground-based space tourism (Museums, Exhibits,Training Fac.)
2. Flights to 100 km altitude (X-Prize Mission)
3. Orbital Flights
4. Orbital Hotel Stays

This paper concentrates on Option (2),a real space experience with zero-g experience and observation opportunities of the Earth's surface.

A single day trip with 8 orbits between 500 and 1000 km altitude at 60° inclination provides great viewing opportunities on large parts of the blue planet. The duration in this case would be 12 h, and the vehicle can return to the launch site without a larger maneuvering effort ( launch and landing take place in the same orbital plane).

A launch vehicle for space tourism must fulfil a number of requirements which are unprecedented in the launch vehicle design and operations history:

1. Very high reliability and safe abort capability at any time
2. Robust, modular design for easy access, inspection and exchange (maintenance and refurbishment)
3. Automated ground processing with short turnaround time
4. Economic vehicle passenger capacity ( >100, cf.ref.2)
5. Passenger accomodation with seating provisions (for launch and re-entry), viewports and volume for zero-g exercise

The economically desired capacity of 100 passengers with their seating and viewport requirements, plus the volume for zero-g floating exercises (one of the attractions of such a tour) leads to a large passenger cabin volume of ca.4 m3 per person which is twice as much as for a standard aircraft cabin.

For cost reasons a relatively short vehicle development period must be planned relying mostly on existing and proven components and experience. However, in contrast to past expendable vehicles, an extensive flight test and demonstration phase must follow where at least two vehicles without passenger cabin perform a large number of operational missions. Such flights can be used, however, for cargo transportation to different orbits and to the space station. In a further interim phase piloted vehicles with a passenger cabin must demonstrate the operational reliability bfore the actual passenger operations can start (Phase 3).

A number of reusable launch vehicle systems has already been studied which in principle could serve as passenger launch systems:

1. SINGLE-STAGE winged or lifting-body vehicle with vertical launch and horizontal landing ( such as the Lockheed-Martin-Martin"> Venture Star")
2. SINGLE-STAGE BALLISTIC launch vehicle with vertical launch and vertical landing ( such as Konkoh Maru, ref.5, Delta Clipper and the BETA Series).
3. TWO-STAGE WINGED vehicle with vertical launch and rocket propulsion (as studied by the Aerospace Corp. by J. Penn, ref.4)
4. TWO-STAGE WINGED systems with horizontal launch mode using air-breathing propulsion (i.e. the MBB/Dasa SÄENGER Concept (ref.3)

FIG.1 shows these four major options in a scale comparison (same payload). The air-breathing vehicle is the largest due to the large hydrogen tanks required, and it has the highest dry mass but the lowest take-off mass of all four options. The winged SSTO is the second in size and dry mass but requires the highest launch mass. The two-stage concept as twin-vehicle is much smaller, but the combined dry mass is only 15 % lower than for the previous option. The smallest vehicle is the ballistic SSTO which has the lowest dry mass, but a 20% higher launch mass than the two-stage twin vehicle.

The specific pros and cons for the four concepts shown in FIG.1 are dis-cussed in the following:

OPTION A, a single-stage rocket-propelled vehicle with vertical take-off and horizontal landing is the approach selected by NASA for a first RLV demonstrator ( X-34). An SSTO vehicle of this type can be a more conventional wing-body configuration (as the Shuttle Orbiter) or a lifting body concept ( Venture Star") which is the more difficult option. The observed weight growth is a good indication for this fact.

An SSTO vehicle of this type is relatively large due to the Hydrogen fuel required in this case, as well as due to the fact that the vehicle net mass is relatively high because the additional structure mass of the wings and aerodynamic control surfaces plus the complex flight control and power system.

Winged vehicles (as well as lifting body configurations) are inherently heavier and larger than ballistic vehicles.The potential range and (often ignored) sensitivity of the net mass fraction with respect to the vehicle size is depicted in FIG.2 as an assembly of all major reusable vehicle studies in the past. Also the actual values and the trend of expendable vehicles are shown for comparison.

OPTION B, the ballistic reusable SSTO vehicle concept with vertical landing has been demonstrated by the experimental Delta Clipper DC-X Project. Since the net mass of such vehicles is lower than for winged or lifting body vehicles both the vehicle size and the development cost are lower than for the other options.

The unconventional landing mode causes somertimes concern, however, the vertical landing capability with landing legs allows in principle the anytime abort landing capability - which does not exist, by example, for winged vehicles during the initial vertical ascent phase.

The third advantage with respect to a space tourism vehicle is the large payload/ passenger cabin volume available on top of the vehicle without special c.g. control problems in case of rear re-entry mode.

OPTION C, the two-stage vehicle option allows the use of Kerosene instead of liquid Hydrogen as fuel. This means an essential reduction of propellant cost and Cost per Flight" (CpF) as shown by J. Penn ( ref. 4 ). However, two different vehicles have to be developed. This is more ex-pensive even though the two stages are relatively small in comparison to a single-stage vehicle with Hydrogen. However, the fuel cost advantage shrinks essentially if one does not use the preset price of liquid Hydrogen, but the reduced cost which would apply in case of large-scale production as shown by the chemical industry (ref.6).

FIG.1 shows a special TSTO-concept, the Twin Configuration where first and second stages externally look alike although their interior is quite different as well as the number of engines. Other configurations with different stage sizes have also been conceived

OPTION D is the most convenient vehicle type from the passenger viewpoint since it resembles mostly the actual aircraft operations. With air-breathing propulsion in the first stage the cruise capability allows great flexibility with respect to launch site avail-ability and mission operations.

However, the large passenger cabin leads to a large second-stage vehicle, the integration of which with the first stage is difficult regarding the aero-dynamic requirements for a hypersonic vehicle in the atmosphere.

The SAENGER Concept with a passenger-configured upper stage (with aircraft seating) allowed some 40 passengers at a launch mass of 400 Mg. For 100 passengers the design would go in size and technology beyond present experience and technology. In addition, although being the most flexible option from the operational viewpoint, it is also the most expensive version regarding development cost. This fact may well exclude this concept from a commercially financed venture ( the first stage vehicle is comparable to a Mach 4 passenger aircraft !).

4. Vehicle Example: BETA-STV-100

4.1 BETA Concept History

The BETA I System Study performed by the author at MBB in 1969/70 under contract from the BMFT (Ministry for Research and Tech-nology) was the first study in Europe dealing with a ballistic single-stage fully reusable launch vehicle. It was a real feasibility study", i.e. the task was to find out whether an SSTO Vehicle would become feasible in the foreseeable future. The answer was yes".

Subsequently, different technical studies were performed on larger-size versions ( BETA II, III, IV) in the 1986 to 1996 period (see TABLE I). BETA IIA was defined in an ESA/ESTEC Study for investigations on the potential performance increase by a variable mixture ratio.

TABLE I : BETA DESIGN HISTORY

Code	Origin	Payload*	GLOW
		Mg	Mg
BETA 1	1970	2	130
BETA IA	1995	6	290
BETA II	1986	15	460
BETA IIrev.	1992	12	600
BETA IIA	1993	18	600
BETA III	1996	20 ( ISS)	800
BETA IV	1997	100	2000

* 90/200 km Orbit, Kourou Launch, 6° Inclin.

All BETA Versions have the same basic features:

• Rear re-entry ( no atmospheric turn maneuver),
• Plug-cluster engine assembly,
• Plug nozzle/heat shield combination,
• Vertical landing on extendible legs (5).

The other basic option is front re-entry plus a rotation maneuver before the final landing which was the McDonnell Douglas Delta Clipper (DC-X) approach. This option allows a greater cross-range capability by the use of aerodynamic control surfaces but it complicates the flight mechanics and the thermal protection system. It also has a c.g.-problem (stability during re-entry) which requires positioning of the payload in the center of the vehicle (between the tanks).

For the BETA Concept there is no c.g.-problem and the payload can be placed on top of the vehicle which is more practical, especially if a second stage must be added for high-energy missions.

FIG.3: Basic BETA SSTO Concept

FIG.3 shows the basic BETA Configuration which is a modular straight-forward and simple design. The only new and more complex feature is the plug-cluster engine configuration with the integrated heat shield.

4.3 BETA-Tourist Launch Vehicle

A BETA-type launch vehicle with a 100 passenger cabin has a launch mass of about 780 Mg (metric tons). 685 Mg propellants are required for ascent leaving a net mass of 75 Mg, including 12.5 Mg propellants for maneuvering, reserves and landing. The total cabin mass is estimated to be 11.6 Mg, plus 8.4 Mg for the passengers and crew. The cargo capability of the unmanned version would be some 17 Mg.

The cabin design is depicted in FIG.4b: a pressure vessel with three levels, each one equipped with 34 seats in a circular arrangement, providing optimum viewing opportunity for each passenger. In the center is a relatively large cylindrical zero-g exercise volume. Ample space for galleys, toilets, storage boxes and equipment is provided. On top of the passenger cabin is a cockpit with seats for the Commander and the Tourist Guide. Although the mission would be performed completely automatic a chief pilot is probably required for psychological reasons. In addition, three hostesses take care of the passengers. The seats can be inclined almost horizontally for the launch and landing phases. The cabin diameter is about 6.5 m.

An overall vehicle mass summary is provided in TABLE II, together with the subsystem mass allocations.

TABLE II : BETA STV-100 Mass Summary

Launch Mass ( GLOW)		780 Mg
Propellant Mass (ascent)		685 Mg
Passenger Cabin (equipped)		11.6 Mg
100 Passengers + 5 Crew		8.4 Mg
Vehicle Dry Mass		62.2 Mg
	Structure	17.0 Mg
	Tanks and Insulation	12.6 Mg
	TPS (Thermal protecion)	5.8 Mg
	Main engine system	16.8 Mg
	OMS/RCS	1.6 Mg
	LGS (Landing Gear System)	2.4 Mg
	Equipment & Margin	6.0 Mg
On-board propellants		12.8 Mg
	OMS/RCS propellants	1.5 Mg
	Residuals, reserve	2.8 Mg
	Landing propellants	8.5 Mg
Vehicle NET MASS		75.0 Mg
	Net Mass Fraction (NMF)	10.9 %

The total thrust level at take-off is 10700 kN in order to provide a launch acceleration of 1.4 g which is optimum for SSTO vehicles (resulting in the minimum delta-V requirement). The number of engines can vary between 12 and 24 units. There is an additional center engine to minimize base drag at launch and to perform the orbital maneuvers (injection, retro impulses). The average specific impulse including the plug nozzle effect at ascent is assumed to be 428 sec (4200 kNs/kg) , with 350 sec at launch and 455 s in vacuum.

During ascent the thrust level needs to be reduced essentially by throttling and/or selected engines' cut-off in order to achieve the maximum performance as well as to limit the thrust acceleration to some 3.5 g with respect to the passengers. Several thrust modulation programs have been tested (FIG.6), as well as thrust vectoring programs, resulting in a remarkable reduction of velocity losses and, therefore, in a reduction of the total required delta-V. Of the three profiles shown in FIG.6 option b with a thrust reduction during the max. Q-conditions showed the best result.

FIG.6: SSTO Thrust Level Profiles optimized with the ALTOS Program

1. S Abitzsch, October 1997, " Economical Feasibility of Space Tourism - A Global Market Scenario", Preprint IAA-97-IAA.1.2.01, IAF-Congress Turin
2. D E Koelle, 1997, " Requirements for Space Tourism Launch Vehicles", Preprint IAA-97-IAA1.2.05, 48th IAF Congress Turin
3. D E Koelle and H Kuczera, October 1990, " SAENGER Space Transportation System", 41st IAF Congress Dresden, Preprint No. IAF-90-175
4. J P Penn and C A Lindley, 1997, " Requirements and Approach for Space Tourism Launch Systems", Preprint IAA-97-IAA1.2.08, 48th IAF-Congress Turin
5. K Isozaki et al, 1994, "Vehicle Design for Space Tourism", Journal of Space Technology and Science (Tokyo), Vol.10, No.2
6. D E Koelle, February 1997, " TRANSCOST 6.1 - Statistical Analytical Model for Cost Estimation and Economical Optimization of Space Transportation Systems", Report TCS-TR-140A(97)
7. N Anfimov, H Kuczera et al, April 1998, " ORYOL-FESTIP Cooperation: Comparison of Concepts and First Conclusions", Paper No. AIAA 89-1544, 8th International Space Planes and Hypersonic Systems and Technology Conference, Norfolk
8. D E Koelle, May 1970, "BETA, A Single-Stage Reusable Ballistic Space Shuttle Concept", 21st IAF Congress, Konstanz/Germany, October 1970, Spaceflight, May 1970
9. D E Koelle, October 1978, " Performance and Cost Analysis for an SSTO+OTV Heavy Cargo Transportation System to GEO", Paper IAF-78A-27, IAF Congres Dubrovnik
10. D E Koelle and W Kleinau, 1986, " The Single-Stage Reusable Ballistic Launcher Concept for Economic Cargo Transportation", Preprint IAF-86-122, Innsbruck, Austria
11. D E Koelle, October 1997, " Cost Engineering - The New Paradigm for Launch Vehicle Design", Preprint IAA-97-IAA.1.04, 48th IAF Congress October 1997 Turin, Italy
12. H Immich and D E Koelle, 1993, " ESA/ESTEC Study on Advanced Rocket Propulsion Technologies"