A cost-engineered launch vehicle for space tourism

Acta Astmnaurica Vol. 44. Nos. 7-12, pp. 649-656. 1999

Pergamon 0 1999 Elsevier Science Ltd. All rights reserved Printed in Great Britain

PII: SOO94-5765(99)00095-S 0094-5765199 $ _ see front matter

A COST-ENGINEERED LAUNCH VEHICLE FOR SPACE TOURISM Dr.-kg. Dietrich E. KOELLE

TCS-TransCastSystems 85521 Ottobrunn / Germany Liebigweg 10

Phone and Fax: ++49-89-6091677 e-mail : [email protected]

ABSTRACT: me paper starts with a set of major requirements for a space tourism vehicle and discusses major vehicle Options PrOpOSed for

mk purpose. It seems mat me requirements can be met best with a Ballistic SST0 Vehicle which has the additional advantage of lOWeSt

development cost compared to other launch vehicle options - important for a commercial development venture. The BETA Ballistic Reusable Vehicle Concept Is characterized by the plug nozzle cluster engine configuration where the plug nozzle serves also as base plate and re-entry heat shield. In this case no athmospheric turn maneuver is required ( as in case-of the front- entry Delta-Clipper DC-Y concept). In our specific ease for space tourism this mode has me avantage that me forces at launch and re- entry are in exactly the same direction, easing passenger seating arrangements. The second basic advantage is the large available volume on top of the vehicle providing ample space for passenger accomodation, visibility and volume for zero-g experience (free floating), one of the major passenger mission requirements. An adequate passenger cabin design for 1 DO passengers is presented, as well as the modem BETA-STV Concept with its mass abXiiOl’IS. Q 1999 Elsevier Science Ltd. All rights

reserved.

1. INTRODUCTON

Space Tourism is the only foreseeable market area where a lar demand can be expecte 8

e and growing 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 tlights 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 economic establishment and operation of an orbital hotel facility.

2. MISSION AND VEHICLE REQUIRElKENTS

2.1 Mission Options

Several options for space tourism can be deflned: (0) Ground-based space tourism

(Museums, Exhibits,Training Fat.) (1) Flights to 100 km altitude

(X-Prize Mission) (2) Orbital Flights (3) Orbital Hotel Stays

This paper concentrates on Option (2l.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

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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).

2.2 Launch Vehicle Requirements

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 desi for easy access, inspection an !? 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 $ned relying mostly on existin

proven components an % 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 passen er cabin must demonstrate B t e operational reliability bfore the actual passenger operations can start (Phase 3).

3. LAUNCH SYSTEM OPTIONS

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

(A) SINGLE-STAGE winged or lifting- body vehicle with vertical launch and horizontal landin ( such as the Lockheed-Martin, Q enture St&‘)

(B) SINGLE-STAGE BALLISTIC launch vehicle with vertical launch and vert.icaI landin

gl ( such as Konkoh

Maru, ref.5, De ta Clipper and the BETA Series).

(C) TWO-STAGE WINGED vehicle with vertical launch and rocket propulsion (as studied by the Aerospace Corp. by J. Penn, ref.4)

(D) TWO-STAGE WINGED systems with horizontal launch mode using air-breathin propulsion (i.e. the MBB/Dasa !z AENGER 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 lar e hydrogen tanks required, and it a as the hi hest dry mass but the lowest take-o d mass of all four options. The winged SST0 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 SST0 which has the lowest dry mass, but a 20% higher launch mass than the two-stage twin vehicle.

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A B

FIG1 : Major Launch Vehicle Options

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 SST0 vehicle of this type can be a more conventfonal win (as the Shuttle %

-body configuration rbiter) or a lifting

body concept (“Venture Star”) which is the more difficult option. The observed weight growth is a good indicatfon for this fact.

An SST0 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 ilight control and power system.

Winged vehicles (as well as lifting

C D

in Realistic Size Comparison

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 SST0 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 le s allows in principle the anytime ?I a ort landing capability - which does not exist, by example, for

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__

3

T

__

. -

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FIG.2: Net Mass Fraction of Winged and Ballistic Vehicles vs. Vehicle Size (PropMass)

winged vehicles during the initial vertical ascent phase. The third advanta

!h e with respect to a

space tourism ve icle 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 Xost per FlighY (CpF) as shown by J. Penn ( ref. 4 ). However, two different vehicles have to be developed. This is more expensive even though the two stages are relatively small in comparison to a single-stage vehicle with Hydrogen.

CONCEPT D is the most con- venient vehicle type from the passenger viewpoint since it re- sembles mostly the actual aircraft operations. With air- breathin stage a

propulsion in the first t e cruise capability

allows great flexibility with respect to launch site availability 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 passen

B erconflljured upper stage (with

aircra t ‘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 regardin

% development cost. This fact may we1 exclude this concept from a commercially financed venture ( the first stage vehicle is comparable to a Mach 4 passenger aircraft !).

However, the fuel cost advantage shrinks essentially if one does not use the present 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 Con6 uration where first and secon % stages externally look alike although their interior is quite different as well as the number of en ines. Other configurations wi $1 different sta e sizes have also been conceive % .


4. Vehicle Example: BETA-WV-100 -- Vertical landing on extendible legs (5).

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 deahng with a ballistic single-stage fully reusable launch vehicle. It was a real ,feasibflity study’, i.e. the task was to find out whether an SST0 Vehicle would become feasible in the foreseeable future. The answer was ,yes’.

The other basic option is front re- entry plus a rotation maneuver before the final landin which was the McDonnell Dou P (DC-X) approach. %

as Delta Clipper his 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 positionin

a the center of the ve of the payload in icle (between the

tanks). 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 de6ned in an ESA/ESTEC Study for investigations on the potential performance increase by a variable nxixture ratio.

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.

TABLE I : BETA DESIGN HISTORY 1

Code Origin Payload* GLOM

Mg Mg BETA I 1970 130 BETA IA 1995 62 290

BETA II 1986 15 460 BETA Ilrev. 1 9 92 12 460 BEi-A IIA 1993 18 600

BETA Ill 1996 20 (ISS) 800

BETA IV 1997 100 2000

FIG.3 shows the basic BETA Configuration which is a modular

W/200 km Orbit, Kourou Launch, 6’ h-&n.

4.2 Basic Design Features

All BETA Versions have the same basic features: __

_*

__

Rear re-entry ( no atmospheric turn maneuver), Plug-cluster engine assembly,

Plug nozzle/heat shield combination, FIG. 3: Basic BETA SST0 Concept


straight-forward and simple desi n. f The only new and more camp ex

feature is the confl

P &

lug-cluster engine ration wi

shiel . the integrated heat

4.3 BETA-Tourist Launch Vehicle

A BETA-type launch vehicle with a 100 passenger cabin has a launch

FIG.4: Cabin Design for 100 Passenger6

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 maneuverin ,

% The total ca reserves and landing.

in mass is estimated to be 11.6 Mg, plus 8.4 M for the passengers and crew. a T e cargo capability of the unmanned version would be some 17 Mg. The cabin design is depicted in FIG.4: a pressure vessel with three levels, each one equipped with 34 seats in a circular arrangement, providing opthnum viewing opportunity for each passenger. In the center is a relatlvely large cylindrical zero-g exercise volume. Ample space for galleys, toilets, sto

3 e boxes and equipment

is provided. n top of the passen er 6 cabin Is a. cockpit with seats for t e

Commander and the Tourist Guide. Although the mission would be performed complete1 automatic a chief pilot is probab y required for 3 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.

The total thrust level at take-off is 1 WOO kN in order to provide a launch acceleration of 1.4 g which is optimum for SST0 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 set (4200 kNs/kg) , with 350 set 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


TABLE II : BETA WV-100 Mass Summary

Launch Mass (GLOW) Propellant Mass (ascent) Passenger Cabin (equipped) 100 Passengers + 5 Crew

Vehicle Dry Mass

780 Mg 685 Mg

11.6 Mg 8.4 Mg

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 -- OMWRCS propellants -- Residuals, reserve -- Landing propellants

12.8 Mg 1.5 Mg 2.8 Mg 8.5 Mg

Vehicle NET MASS 75.0 Mg -- Net Mass Fraction (NMF) 10.9 %

FIG.5 : Plug Cluster Engine Concept

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.G), 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.

IIY ( bl

FIG. 6: SST0 Thrust Level Profiles optimized with the ALTOS Program

5. REFERENCES

1. S.ABITZSCH: Economical Feasibility of Space Tourism - A Globa Market Scenario, heprint IAA-97-h&%. 1.2.01. IAF-congress Turin. Oct. 1997

2. D.E.KOELLE: Requirements for Space Tourism Launch Vehicles, Preprint WA-97 IAA1.2.05.43th IAF Congress Turin.1997

3. D.E.KOELLE. H.KUCZERA: SAJZNGER Space Transportation System, 41st IAF- Congress Dresden, Oct. 1990. Preprint No. IAF-go- 175


4.

5.

6.

7.

J.P.PENN and C.A.LINDLEYz Requirements 8. D.E.KOELLE: BETA, A Single-Stage and Approach for Space Tourism Launch Reusable Ballistic Space Shuffle Concept, Systems, Preprint IAA-97-IAA1.2.08.48th 21st IAF Congress, Konstanz/Gennany. IAF-Congress Turin. 1997 Oct. 1970. Spaceflight. May 1970

KISOZAKI et al.: Vehicle Design for Space Tourism, Journal of Space Technology and Science (Tokyo], Vol.10, No.2. 1994

9. D.E.KOELLE: Performance and Cost Analysis for an SSTO+OTV Heavy Cargo Transportation System to GEO. Paper IAF- 78A-27. IAF Congres Dubrovnik, Oct. 1978

D.E.KOELLE: TRANSCOST 6.1 - Statistical Analytical Model for Cost Estimation and Economical Optimtzation of Space Trans- portation Systems, Feb. 1997. Report TCS- TR- 14oA(97)

lO.D.E.KOELLE and WKLEINAU: The Slngle- Stage Reusable Ballistic Launcher Concept for Economic Cargo Transportation. Preprint IAF-86- 122, Innsbruck, Austria

N.ANFlMOV. HXUCZERA et al.: ORYOL- FESTIP Cooperationz Comparison of Concepts and First Conclusions. Paper No. AIAA 89- 1544,8th International Space Planes and Hypersonic Systems and Technology Conference, Norfolk, April 1998

11. D.E.KOELLE: Cost Engineering - The New Paradigm for Launch Vehicle Design, Pre- print TAA-97-IAA. 1_04,48th TAF Congress Oct. 1997 Turin. Italy

12. H.IMMICH. D.E.KOELLE: ESA/ESTEC Study on Advanced Rocket Propulsion Technologies, 1993

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A cost-engineered launch vehicle for space tourism