Rockwell Star-Raker SSTO concept to support SPS (1978)

Bibliography/References:

Star-Raker: An Airbreather/Rocket-Powered, Horizontal Takeoff Tridelta Flying Wing, Single-Stage-to-Orbit Transportation System – SSD 79-0082 (7.12 MB PDF) (Original 18 MB PDF)
Earth-to-LEO Transportation System for SPS – IRD Data Sheet 243 (14.3 MB PDF)
NASA CR 3321: Satellite Power Systems (SPS) Concept Definition Study – Volume IV: Transportation Analysis Excerpt (1~ MB PDF)
NASA TM 58238: Satellite Power System: Concept Development and Evaluation Program – Volume VII: Space Transportation Excerpt (610 KB PDF)

Description

The North American Rockwell Star-Raker was an outgrowth of the late 1970s Satellite Power System (SPS) studies; which converged around building sixty SPS satellites over a period of thirty years.

Each satellite would mass 35,000~ metric tonnes, requiring 70,000 metric tons to be delivered to LEO each year. With a payload of 91 metric tons per launch vehicle, 770 flights per year, or 2.1 flights per day would be required to meet the build requirements.

The size of this program drove the need for long term routine operations analogous to commercial airline/airfreight operations. While “conventional” Super-Heavy Launch Vehicles with 400 metric ton payloads could reduce the launch rate requirement each year to just 175 flights, the SHLV concept (water recovery of stages with subsequent refurbishment) faced severe scheduling constraints due to refurbishment, stacking and launch pad cycle times.

Thus, Rockwell continued with previous studies into advanced horizontal takeoff, single-stage-to-orbit (HTO-SSTO) aerospace vehicles due to previous studies showing significant operational advantages from said vehicles.

One of the configurations that emerged became known as the “Star-Raker”.

Star-Raker Bulletpoints

Operational/Handling Features

• Total Vehicle Re-usability with Rapid Turnaround.

• Ferry capability between airfields (with cargo if necessary).

• Requires runways 8,000 to 14,000 ft long (2,440 to 4,270 m), with a takeoff velocity of 225 knots or greater and landing velocities of about 115 knots.

• Ability to reach any LEO plane from alternate launch sites (KSC, Vandenberg, etc) and return to the same site (including single orbit passes).

• 200,000 lb (90,718 kg) maximum payload capacity in a 20 x 20 x 141.5 ft cargo bay (56,600 ft3) accessed via hinged nose.

• Delivery of 196,600 lbs (88,904 kg) to a 300 nautical mile @ 28.5 deg inclination orbit flying directly from Kennedy Space Center for a cost of $10 to $15 per pound (1978 dollars).

• Cargo bay similar to C-5A Galaxy configuration and supports the use of MATS and Airlog cargo handling systems.

Propulsion Features

• Ten hydrogen fueled high bypass supersonic turbofan/air-turbo-exchanger/ramjet engines, each with 140,000 lbf of thrust. Engine design to be based off the GE CJ805 axial flow turbojet, P&W SWAT 201 turbofanramjet, Aerojet Air Turborocket, Marquardt Variable Plug Nozzles, ramjet engines, and Rocketdyne tubular cooled rocket engines. Subsonic ISP of 9,700 seconds that linearly reduces to 4,000 seconds at 366 m/sec. Supersonic ISP that reduces from 4,000 seconds at 366 m/sec to 3,500 seconds at 1,700 m/sec.

• Three hydrogen fuelled rocket engines, each with 1.06 million lbf of thrust and an ISP of 455 seconds.

Flight Profile

A typical profile flown from Kennedy Space Center to a 300 n.mi orbit at 28.5 deg inclination was to be:

1. Runway takeoff under high-pass turbofan/airturbo exchanger (ATE)/ramjet power, with the ramjets acting as supercharged afterburners.

2. Jettison and parachute recovery of landing gear used only for launch.

3. Climb to optimum cruise altitude with turbofan power.

4. Cruise at optimum altitude, Mach number, and direction vector to earth's equatorial plane, using turbofan power.

5. Execute a large-radius turn into the equatorial plane with turbofan power.

6. Climb subsonically at optimum climb angle and velocity to an optimum altitude, using high bypass turbofan/ATE/ramjet (supercharged afterburner) power.

7. Perform an optimum pitch-over into a nearly constant-energy (shallow Y_angle) dive if necessary, and accelerate through the transonic region to approximately Mach 1.2, using turbofan/ATE/ramjet (supercharged afterburner) power.

8. Execute a long-radius optimum pitch-up to an optimum supersonic climb flight path, using turbofan/ATE/ramjet power.

9. Climb to approximately 29 km (95 kft) altitude, and 1900 m/s (6200 fps) velocity, at optimum flight path angle and velocity, using proportional fuel-flow throttling from turbofan/ATE/ramjet, or full ramjet, as required to maximize total energy acquired per unit mass of fuel consumed as function of velocity and altitude.

10. Ignite rocket engines to full required thrust level at 6200 fps and parallel burn with airbreathing engines to 7200 fps.

11. Shut down airbreather engines while closing airbreather inlet ramps.

12. Continue rocket power at full thrust.

13. Insert into an equatorial elliptical orbit 91 x 556 km (50 x 300 nmi) along an optimum lift/drag/thrust flight profile.

14. Shut down rocket engines and execute a Hohmann transfer to 556 km (300 nmi).

15. Circularize Hohmann transfer.

16. Release Payload or dock with Space Station at that orbit.

17. Perform delta-v maneuver and insert into an equatorial elliptical orbit 91 x 556 km (50 x 300 nmi) in preparation for re-entry.

18. Perform a low-gamma (flight path angle), high-alpha (angle of attack) re-entry deceleration profile very similar to Space Shuttle to approximately Mach 6.

19. Reduce alpha (angle of attack) to appropriate angle for maximum lift/drag ratio for high speed glide and cross range maneuvers to subsonic velocity (Mach 0.85).

20. Open inlets and start some airbreather engines.

21. Perform powered flight to landing field, land on runway, and taxi to jetway. Flyback fuel requirements include approximately 300 nmi subsonic cruise and two landing approach maneuvers (first approach waveoff with fly-around for second approach).