In fact, the progress in propulsion by high-pressure LOX/ LH2 engines with 460 sec Isp_vac is the key for the realization of a single-stage shuttle. As an example, these engines if used with the existing Saturn S-IVB stage, would orbit the vehicle as it is, with zero payload. In combination with the S-II Stage (450 tons launch weight) even a 12 ton payload would be feasible. These stages represent the technology of 1960/65. With the technology of 1970/75, an improvement certainly seems possible) namely the recovery of such a stage by applying Apollo/LM techniques.

A certain problem of single stage shuttles is the limitation to low orbits. For higher velocity requirements the performance becomes uneconomic, and a second stage is required. Otherwise, recent trends in space transportation led to the Shuttle/Tug Concept, where the task of transportation in space (orbit to orbit or orbit to Moon) is conducted by a re-usable space tug.

In this case, the shuttle target orbit can have a low altitude and a low inclination, decoupling it from the Space Station Orbit of 550/500 km. The latter is one of several missions and it is no good place for continuing to geostationary orbits, lunar or interplanetary orbits for energy reasons as well as for operational reasons.

The logical consequence of a shuttle/tug transportation system will lead to a special type of "Space Operations Base" or Space Service Station (SSS) which is different from the present Space Station for space research and application tasks.

The SSS role can be defined as follows:

1. Earth-to space shuttle terminal
2. Space-tug operations base
3. Payload-tug integration and checkout platform
4. Vehicle assembly and checkout basis for larger vehicles requiring more than one shuttle flight
5. Tank farm (propellant storage facility).

Fig. 1 illustrates this operations concept, which has the performance advantage that - in case of the Cape Kennedy Launch Base - the SSS can be placed in a 28.50 degree orbit (compared to the 55 degree Space Station Orbit) at a lower altitude (appx. 200 NM or 370 km). The payload for this proposed SSS-orbit is more than twice as large as for the 550/500 km orbit, using the single-stage shuttle, or appx. 50% higher in case of the two-stage winged shuttle.

BETA stands for "Ballistisches Einstufiges Träger-Aggregat" (Ballistic Single-stage Carrier Vehicle) and it represents a combination of features of present launch vehicles, the Apollo Capsule and the LM Lunar Lander: re-entry technique and technology as in case of Apollo, final landing approach and touchdown as in case of the lunar module.

BETA is an unmanned, fully automatic shuttle system. It does not require a crew since no manual fly-back capability is required. Otherwise, a crew can be added if desired for certain orbital operations.

The BETA concept is depicted in figs. 2 and 3. It is characterized by the following features:

1. A short conical body (small length/diameter ratio, low c.g.) with heat-shield for re-entry
2. Use of the heat-shield as a plug-nozzle for performance increase
3. A propulsion system consisting of 12 or more single high-pressure LH2/ LOX engines arranged around the central plug-nozzle (heat-shield)
4. 6 retractable legs for the final vertical landing phase.

This concept has been selected as a baseline configuration out of a variety of configurations shown in fig. 4:

BETA has a single large hydrogen tank and a toroidal (multi-cell) LOX tank in a conical arrangement. This leads to a low c.g. for good stability in flight and on ground. The large heat shield serves during ascent as a plug-nozzle for performance improvement of a series of high-pressure rocket motors. These motors are of conventional configuration and have a small thrust level compared to the vehicle size. They are rigidly mounted around the LOX tank as shown in fig. 5. No gimbaling is required since thrust vector control can be achieved by single engine throttling. The relatively large bottom area leads to small aerothermodynamic loads (1/3 of Apollo) during re-entry and high engine performance during ascent. Final landing will be performed by retro-thrust (4 motors only) and six retractable legs. The payload is mounted atop of the vehicle with or without a special fairing. This arrangement means much less geometrical restrictions than a cargo compartment within a winged shuttle.

Items	Mass (kg)
LH2 tank (Ti-alloy A1 V64T)	600
...insulation, fittings, etc.	380
LOX tank (Al-alloy, AlZnMg3)
...incl. Insulation, fittings, etc.	1,550
Helium pressure vessels (within LOX tank)	420
Propellant flow system (lines, valves, etc.)	655
Rocket engines (12)	2,200
Thrust frame and engine mounting	660
Payload and fairing adapter ring	162
Upper structure	140
Lower main structure (Ti-alloy)	770
Plug nozzle/head shield	925
Tanks for return flight propellants	150
Landing legs (6)	550
Auxiliary propulsion system for attitude control	105
Central computer and sequencer	45
Reference system with electronics	40
Electrical power supply and harness	140
Telemetry system and sensors	54
Safety system	30
Landing radar equipment	90
Other items	80
Margin	254
Dry Weight	10,000
Propellant residuals	1,300
Propellants for attitude control	250
Return flight propellants	950
Nominal propellant weight for ascent	115,000
Total launch weight without payload	127,500
Payload	4,000
Total launch weight	131,500

Table 1 shows a weight break-down for the minimum size BETA I vehicle which has been analyzed in more detail since it is the most critical case. The table shows rather conventional material assumptions and weight estimates, which still allow some improvements.

The principal advantage of the single-stage BETA-Vehicle is the possibility to launch it from places in Europe, since no stages or parts fall away during ascent. Vertical take-off and vertical landing in combination with the landing leg system allows continuous abort capability in the critical launch phase. It also allows an outstanding means of test operations, which continuously lead from ground testing to flight tests.

The ascent trajectory optimization of a single-stage vehicle of the BETA type is more complex than for a multistage vehicle: this is due to the fact that beside the atmospheric and dynamic parameters a thrust optimization has to be made very carefully since it influences the performance to a large extent.

The result is first that the launch acceleration should be higher than usual: appx. 1.5 g, and that a thrust programme has to be applied which decreases the thrust level down to 10% in a way that the optimum ascent time is achieved. This optimum ascent time (equal to the engine burning time) was found to be 500 sec for 200 km orbit, as shown in fig. 6. This optimization process showed no higher acceleration than 3.5 g.

The propellant mixture ratio was also involved in the optimization: it is shifting from 5.5 at launch to 8.0 before cutoff, taking into account the large difference in expansion ratio.

BETA remains in its orbit as required by operations and flight mechanics before a braking impulse (given by 4 engines) initiates return to Earth. A secondary propulsion system for orbital maneuvers, attitude control and propellant orientation has aligned the vehicle before and allows control during ascent.

The plug-nozzle heat-shield protects the vehicle during the critical re-entry phase, while the engine nozzles are cooled down by hydrogen circulation. Finally, a constant descent velocity of appx. 70 m/sec is reached (depending on the ballistic factor) which has to be reduced to zero by a final brake impulse given by 4 engines. The landing shock is absorbed by the legs.

Fig. 7 shows the major descent phases. The LM Lunar Landing Vehicle has demonstrated point landing capability on unprepared terrain and this method should be feasible on Earth too. The trajectory has to be corrected such that BETA reaches a point 5 to 10 km above its landing area and performs practically a vertical descent with a limited cross-range capability.

The performance analysis chapter deals with the fundamentals of a single-stage ballistic vehicle, which are the engine system performance and the structural efficiency as well as the velocity requirement for the mission range envisaged. Closely connected with this analysis is the problem of effective payload and sizing of the vehicle.

The vacuum performance of a single high-pressure rocket engine with a nozzle area ratio of about 250 is about 460 sec Isp. This is the result of design studies of Rocketdyne, Aerojet and Pratt & Whitney for the US Space Shuttle engine. Further, it has been proven that chamber pressures well above 200 atm can be realized, by testing rocket chambers built by MBB in the United States (Reno Test Site). The topping cycle principle applied for these engines has been developed and demonstrated by MBB with a 6 ton LOX/Kerosene engine.

The problem of rocket engines with a conventional (bell) nozzle is the fact that they can be optimized only for one altitude, which means losses in all other flight regions. This is shown in fig. 8: the larger the nozzle and so higher the expansion ratio, so better the specific impulse in space but more worse after launch. The latter means also larger and heavier engines to achieve the required launch acceleration.

The plug nozzle is a solution to this problem since the expansion is self-adapting to the external pressure and, therefore, delivering always the optimum specific impulse. On the other hand, the development of a plug-nozzle engine with 10 m diameter and several hundred tons thrust represents a relatively large development effort, even if the segmented design concept is used.

As a compromise, the BETA concept assumes a series of 12 to 20 single rocket engines with a small thrust level compared to the vehicle mass, grouped around a central plug which serves also as the heat-shield during re-entry. This means that in the initial launch phase only an area ratio of 1:35 is effective, resulting in an Isp of 380 sec (fig. 8). Approximately from an altitude of 10 km the plug nozzle becomes effective and allows the increase of the expansion ratio to about 1:500 in more than 100km altitude with an Isp of about 470 sec. The average specific impulse which can be calculated using the ascent profile is about 458 sec, since only a relatively small part of the trajectory is within the atmosphere.

Beside the engine performance the structural efficiency, or better the net weight, is responsible for the payload which can be achieved for a certain mission. For the present analysis the socalled "design factor" had been used which is the ratio of net weight (or burnout mass without payload) to propellant mass (usable mass).

Having established an average Isp-value a design chart can be prepared for a specific mission, expressed in required delta V, using the design factor as parameter. Fig. 9 shows such a chart for a typical single-stage mission. As examples the S-IVB and S-II stages with their design factors are shown, indicating the influence of the vehicle size (in terms of propellant mass) on the design factor. In addition, these examples are reference values for the technology standard of 1960/65. This standard, indicated by the dashed line in fig. 9, will certainly move upward with time. Also indicated is Phil Bono's SASSTO concept with a very advanced technology. The latter is not really a requirement for single-stage vehicles, as it is thought widely, since a relatively small increase in size (or propellant mass) allows a more standard technology for the same payload. This can be seen clearly from fig. 9. The diagram is only valid for the ascent phase or non-reusable vehicles. For the final launch weight estimate the additional weight for the return propellants and the required equipment has to be taken into account. The BETA I vehicle with 115 t propellant mass as a minimum size solution and a design factor of 0.10 has according to fig. 9 a payload of 4 tons for an expendable mission (no return). In case of a return mission the payload is only 2 tons.

BETA II with 10 tons payload as a European Mini-Shuttle requires appx. 350 tons launch mass and a BETA III vehicle with the US-Space Shuttle payload size of 20 tons some 600 tons launch mass. In both cases the additional weight requirement for the return phase has been taken into account.

The mission velocity requirement for single-stage launch vehicles is composed of the items and typical values shown in table 2.

The velocity required for return is small: appx. 3% of the total amount. Therefore, the penalty for return is not very large, taking into account the fact that the heat-shield is used as nozzle structure.

1 Orbital velocity (200 km)	7,314 to	7,774 m/s
2 Gravitation losses		1,350 m/s
3 Drag losses	400 to	800 m/s
4 Orbital maneuvers (assumption)		76 m/s
Ascent	9,140 to	10,000 rn/s
	(equatorial orbits)	(polar orbits)
5 Deorbit impulse	50 to	100 m/s
6 Range dispersion correction	10 to	120 m/s
7 Brake impulse	60 to	80 m/s
8 Reserve for hovering	80 to	100 m/s
Total delta V	9,340 to	10,400 m/s

The flight test facilities and operations required for the BETA Shuttle Concept are certainly considerably less complex and less costly than for a two-stage winged system.

Only one system has to be developed and no separation problems have to be solved, flight tests evolve out of ground testing: the critical phase of take-off can be tested repeatedly - a unique feature compared with any other concept.

A major development effort lies in the engine system (not the single engine): the optimization of the plug nozzle and engine integration.

Another problem is the heat-shield which should be light-weight and survive more than one flight to reduce the refurbishment cost. Ablative vs. hydrogen-cooled metal shields have to be investigated.

The development cost for minimum size BETA I Vehicle have been estimated to some 500 MIO $ (European cost level 1969). This includes engine development and fabrication of three vehicles.

A BETA III Vehicle with 20 tons payload capability would require some 2 BILLION $ or less than 1/3 of the presently envisaged development cost of an equivalent two stage winged shuttle.

The specific transportation cost are influenced by

• the vehicle size, resp. payload size
• the refurbishment cost
• the number of launches per year, and
• the possible number of re-uses for the same vehicle.

A preliminary analysis for the small BETA I vehicle led to specific transportation cost of some 300 $/kg, while for the large BETA III vehicle the specific cost should be below 100 $/kg.

1. The development of a single stage ballistic space shuttle is feasible with the present technology.
2. The transportation cost Earth-to-Orbit can be reduced to some 200 $/kg or less.
3. The single-stage concept allows new possibilities for launch ranges since no danger by expendable parts or stages has to be expected. This means that launches from and landings within Europe would be feasible.
4. The BETA Concept seems to be a solution for specific European requirements since
   • there is no manned space flight programme,
   • there does not exist the 500 km/550 space station target orbit, and
   • there is no requirement for a large cross range capability.
5. From the very principle the BETA Concept seems to be the final solution for the space transportation problem since it combines operational simplicity with lowest cost both for development and specific payload cost.

It is worth further investigation.

1. Messerschmitt-Bölkow-Blohm GrnbH, August 1969, " BETA, Ballistisches Einstufiges Träger-Aggregat", Berich UR-V-25(69)
2. Phil Bono, May 1967, " The enigma of booster recovery - ballistic or winged?", SAF Space Technology Conference, Palo Alto, Calif., USA.
3. Messerschmitt-Bölkow-Blohm GmbH, May 1969, " Hochdrucktriebwerke für die Europäische Raumfahrt", Bericht TR-819.
4. G Zietlow, " Systemanalytische Untersuchung von Werkstoff und Bauweisen der Neptun-Aussenstruktur", TUB-IR 1968/16.
5. D Koelle, 1964, " Theorie und Technik der Raumfahrzeuge", Verlag Berliner Union
6. Messerschmitt-Bölkow-Blohm GmhH, March 1969, " Liquid propellant transfer and apogee injection systems", ELDO-Study 16-2
7. H P Riemann, March 1963, 'Vergleich verschiedener Düsenformen für Flüssigkeitstriebwerke', Bericht TR442, ( MBB)
8. Dr. Schmidt, 'Vergleich und Marktanalyse von Zellenbauwerkstoffen für kryogene Treib-stoffe', Bericht TR-590 ( MBB).
9. W Froitzheirn and K Krumm, " Theoretische Untersuchungen zum intakten Wiedereintritt von Radionuklidgeneratoren", Bericht FM-388 ( MBB).
10. Contimet, 'Eigenschaften von Tital und Titan-Legierungen', Katalog
11. AIAA, Fifth Structures and Materials Conference, April 1964
12. Phil Bono et al., " The influence of unconventional structures and advanced materials on booster re-useability", AIAA-Paper