Home - Search - Browse - Alphabetic Index: 0- 1- 2- 3- 4- 5- 6- 7- 8- 9
A- B- C- D- E- F- G- H- I- J- K- L- M- N- O- P- Q- R- S- T- U- V- W- X- Y- Z

---

Beta

---

Beta

AKA: Ballistisches Einstufiges Traeger-Aggregat I;Ballistisches Enistufiges Traeger-Aggregat I. Status: Study 1969. Payload: 2,000 kg (4,400 lb). Thrust: 2,000.00 kN (449,600 lbf). Gross mass: 127,500 kg (281,000 lb). Unfuelled mass: 10,000 kg (22,000 lb). Specific impulse: 470 s. Specific impulse sea level: 380 s. Height: 48.00 m (157.00 ft). Diameter: 7.65 m (25.09 ft). Span: 10.00 m (32.00 ft). Apogee: 90 km (55 mi).

BETA ("Ballistisches Einstufiges Träger-Aggregat" - Ballistic Single-stage Carrier Vehicle) represented a combination of features of existing 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 was an unmanned, fully automatic shuttle system. A crew could be added if desired for certain orbital operations. The following features characterized the BETA concept:

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

Mass breakdown for the BETA I with 115 t nominal propellant mass was as follows:

• LH2 tank (Ti-alloy A1 V64T) 600 kg
  
• LH2 tank insulation, fittings, etc. 380 kg
• LOX tank (Al-alloy, AlZnMg3) including insulation, fittings, etc. 1,550 kg
• Helium pressure vessels (within LOX tank) 420 kg
• Propellant flow system (lines, valves, etc.) 655 kg
• Rocket engines (12) 2,200 kg
• Thrust frame and engine mounting 660 kg
• Payload and fairing adapter ring 162 kg
• Upper structure 140 kg
• Lower main structure (Ti-alloy) 770 kg
• Plug nozzle/head shield 925 kg
• Tanks for return flight propellants 150 kg
• Landing legs (6) 550 kg
• Auxiliary propulsion system for attitude control 105 kg
• Central computer and sequencer 45 kg
• Reference system with electronics 40 kg
• Electrical power supply and harness 140 kg
• Telemetry system and sensors 54 kg
• Safety system 30 kg
• Landing radar equipment 90 kg
• Other items 80 kg
• Margin 254 kg

Propellants

• Propellant residuals 1,300 kg
• Propellants for attitude control 250 kg
• Return flight propellants 950 kg
• Nominal propellant weight for ascent 115,000 kg

Total launch weight without payload = 127,500 kg
Payload = 4,000 kg (one-way mission) or 2,000 kg (recoverable version)
Total launch weight = 131,500 kg

The principal advantage of the single-stage BETA-Vehicle was the possibility to launch it from any location in Europe, since no stages or parts fell away during ascent. Vertical take-off and vertical landing in combination with the landing leg system allowed continuous abort capability in the critical launch phase. It also allowed an outstanding means of test operations - incremental advances from ground testing to flight tests.

The ascent trajectory optimization of a single-stage vehicle of the BETA type was more complex than for a multistage vehicle. The vehicle performance was much more sensitive to thrust optimization taking into account the atmospheric and dynamic parameters. The results showed that the launch acceleration should be higher than that for conventional vehicles: 1.5 g. Thereafter a thrust program had to be applied which decreased the thrust level down to 10% in such a way that the optimum ascent time was achieved. This was found to be 500 sec for a 200 km orbit, with the acceleration never exceeding 3.5 g. The propellant mixture ratio was also involved in the optimization: it was shifted from 5.5 at launch to 8.0 before cutoff, taking into account the large difference in expansion ratio.

BETA would remain in orbit as required by operations and flight mechanics before a braking impulse (given by 4 engines) initiated return to Earth. A secondary propulsion system for orbital maneuvers, attitude control and propellant orientation aligned the vehicle before the maneuver and allowed control during descent. The plug-nozzle heat shield protected the vehicle during the critical re-entry phase, while the engine nozzles were cooled by hydrogen circulation. Finally, a constant descent velocity of about 70 m/sec was reached which was then reduced to zero by a final brake impulse given by 4 engines. The legs absorbed the final landing shock.

The LM Lunar Lander had demonstrated point-landing capability on unprepared terrain and this method should be feasible on Earth too. The trajectory had to be corrected such that BETA reached a point 5 to 10 km above its landing area and performed practically a vertical descent - it had only a limited cross-range capability.

The vacuum performance of a single high-pressure rocket engine with a nozzle area ratio of about 250 was about 460 sec specific impulse, as indicated in contemporary design studies of Rocketdyne, Aerojet and Pratt & Whitney for the US Space Shuttle engine. MBB had demonstrated high-pressure engines with this chamber pressure in the United States at the Reno Test Site. The topping cycle principle applied for these engines was demonstrated by MBB with a 6 metric ton LOX/Kerosene engine.

The problem of rocket engines with a conventional (bell) nozzle was the fact that they could only be optimized for one altitude, which meant losses in all other flight regions. A larger nozzle with a higher expansion ratio provided better specific impulse in space but much lower at launch. It also meant larger and heavier engines to achieve the required launch acceleration. The plug nozzle was a solution to this problem since the expansion was self-adapting to the external pressure and was believed to always deliver optimum specific impulse. However the development of a plug-nozzle engine with 10 m diameter and several hundred metric tons thrust represented a relatively large development effort, even if the segmented design concept was used.

As a compromise, the BETA concept assumed a series of 12 to 20 single rocket engines with a small thrust level compared to the vehicle mass, grouped around a central plug that also served as the heat shield during re-entry. In the launch phase this resulted in an effective area ratio 35:1 and a specific impulse of 380 sec. Above 10 km the plug nozzle became effective, finally attaining an effective expansion ratio of 500:1 and a specific impulse of 470 seconds at 100 km altitude. The average specific impulse using the selected ascent profile was 458 sec.

The additional velocity required for recovery of the BETA was relatively small - 3% of the total amount. Therefore, the penalty for recovery was not very large, taking into account the fact that the heat-shield was used as nozzle structure. The velocity requirements were calculated as follows:

• Orbital velocity (200 km) 7,314 to 7,774 m/s (equatorial/polar orbits)
• Gravitation losses 1,350 m/s
• Drag losses 400 to 800 m/s
• Orbital maneuvers (assumption) 76 m/s
• Total Ascent 9,140 to 10,000 m/s (equatorial/polar orbits) (equatorial orbits) (polar orbits)

• Deorbit impulse 50 to 100 m/s
• Range dispersion correction 10 to 120 m/s
• Brake impulse 60 to 80 m/s
• Reserve for hovering 80 to 100 m/s
• Total delta V 9,340 to 10,400 m/s (equatorial/polar orbits)

The flight test facilities and operations required for the BETA Shuttle Concept were certainly considerably less complex and less costly than for a two-stage winged system. Only one system had to be developed and no separation problems had to be solved. Flight tests could evolve out of ground testing - the critical phase of take-off could be tested repeatedly - a unique feature compared with any other concept. The major development effort would lie in the engine system (not the single engine): the optimization of the plug nozzle and engine integration. Another problem was the heat shield, which had to be lightweight and survive more than one flight to reduce the refurbishment cost. Ablative and hydrogen-cooled metal shields would have to be investigated. The specific transportation cost would be influenced by the vehicle size and resulting 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 a specific transportation cost of some 300 $/kg.

LEO Payload: 2,000 kg (4,400 lb) to a 90 km orbit at 6.00 degrees. Development Cost $: 500.000 million. Launch Price $: 0.600 million in 1969 dollars in 1969 dollars.

Stage Data - Beta

• Stage 1. 1 x Beta. Gross Mass: 450,000 kg (990,000 lb). Empty Mass: 40,000 kg (88,000 lb). Thrust (vac): 5,736.000 kN (1,289,504 lbf). Isp: 460 sec. Burn time: 318 sec. Isp(sl): 409 sec. Diameter: 7.65 m (25.09 ft). Span: 10.00 m (32.00 ft). Length: 40.00 m (131.00 ft). Propellants: Lox/LH2. No Engines: 13. Engine: MBB-ATC500. Status: Study 1969.

Beta II German SSTO VTOVL orbital launch vehicle. Beta II was Dietrich Koelle's nominal 350 metric tons lift-off mass SSTO design for launch of a 10 metric ton European spaceplane.