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DC-X

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DX-XA Flight 2 Credit: NASA

AKA: Clipper Graham;DC-XA;Delta Clipper Experimental;Delta Clipper-Experimental;SX-1. Status: Retired 1996. First Launch: 1993-08-18. Last Launch: 1995-07-07. Number: 8 . Thrust: 223.00 kN (50,132 lbf). Gross mass: 16,320 kg (35,970 lb). Unfuelled mass: 7,200 kg (15,800 lb). Specific impulse: 373 s. Specific impulse sea level: 316 s. Burn time: 127 s. Height: 14.00 m (45.00 ft). Diameter: 3.05 m (10.00 ft). Span: 4.10 m (13.40 ft). Apogee: 3.00 km (1.80 mi).

The DC-X was an experimental vehicle, 1/3 the size of a planned DC-Y vertical-takeoff/vertical-landing, single stage to orbit prototype. It was not designed as an operational vehicle capable of achieving orbital flight. Its purpose was to test the feasibility of both suborbital and orbital reusable launch vehicles using the VTOVL scheme. The DC-X flew in three test series. The first series ran from August 18 to September 30, 1993, before the initial project funding ran out in late October 1993. Additional funding was provided and a second series was conducted June 1994-July 1995.

The Single Stage Rocket Technology program was an SDIO (Space Defense Initiative Office) program to demonstrate technology readiness for an SSTO vehicle. The SDIO (Strategic Defense Initiative Organization) issued a call for proposal for design studies of a single stage to orbit launch vehicle in August 1990. Six companies responded, four proposing vertical takeoff concepts (McDonnell-Douglas, Rockwell, General Dynamics and Boeing) and two proposing air-launched concepts (Grumman and Third Millennium). A total of $ 12 million in study contracts were awarded to the four companies proposing VTO designs, reflecting SDIO's preference for VTOL concepts studied in an Aerospace Corporation report of the previous year.

Douglas and General Dynamics proposed VTOVL designs. Boeing's final design was an improved version of its HTHL sled- or rail-launched RASV (Reusable Aerodynamic Space Vehicle) powered by uprated Shuttle Main Engines. Rockwell proposed a VTHL vehicle, similar in layout to the Space Shuttle Orbiter, but with an aerospike engine.

Phase II of the program would consist of fabrication of an 'X' subscale demonstrator and design of a 'Y' prototype vehicle. Reflecting SDIO's predilections, Douglas was finally awarded the $60 million Phase II contract The Douglas DC-X was designed primarily to validate the turnaround and landing maneuver required for the company's complex nose-first entry concept (as opposed to the base entry approach of prior VTOVL SSTO designs). This approach would, the company believed, provide a wide cross range, always attractive to the Air Force for operational reasons. In commercial applications nose-first entry would reduce the loads experienced by the crew and passengers to under 2G, compared with 3G for base-first lifting entries. The DC-X was specifically not designed to demonstrate the most critical issue for any practical SSTO: structural mass fraction

Under a 2-year, $58-million Phase II contract, MDSSC and its team-mates were using a rapid prototype approach to design and build a one-third-size experimental vehicle the DC-X, and ground support and operations systems which, through a series of suborbital flights, was to:

• Verify vertical takeoff and landing
• Demonstrate subsonic maneuverability
• Validate "airplane-like" supportability/maintainability concepts
• Demonstrate rapid prototyping development approach

The SSRT program consists of three phases. Phase I began in August 1990 and consisted of a $12 million design and risk reduction competition. At that time, the program was focused on multiple single stage-to-orbit concepts which were found to be potentially viable, including vertical take off and landing (VTOL), horizontal takeoff and landing (HTOL) and vertical takeoff and horizontal landing (VTHL) configurations. Following completion of Phase I, a two-year $58 million Phase II contract was competitively awarded to McDonnell Douglas Aerospace, Huntington Beach, Calif., in August 1991. The program was subsequently restructured and focused on building the DC-X and enabling suborbital RLVs for potential use by BMDO. The DC-X design emphasized simplified ground and flight operations, vehicle maintenance, rapid turnaround, and operational characteristics that were also relevant to future orbital vehicles. For example, the highly automated control center for this system was manned by only three people: two for flight operations and one for ground operations and servicing. Successful completion of the DC-X testing in mid 1993 was to form the basis for a Phase III "go/no go" decision by the Department of Defense to develop a follow-on Advanced Technology Demonstrator for support of flight tests and experiments. If a decision was made to proceed with Phase III, the program was to be transferred to another agency.

A systems ground test facility was activated at NASA's White Sands Test Facility (WSTF), and a launch and recovery site at WSMR. The completed DC-X vehicle was tested on a modified propulsion test stand at the WSTF prior to flight testing at WSMR. Maintenance and ground support techniques required for the flight test phase of the program was to be tested, evaluated and refined at WSTF. The aircraft-like flight test program was to began with low altitude hover flights, gradually increasing in altitude and duration, and was to lead to suborbital flights to approximately 5500 m.

Funding was stop-and-go however. By July 1995 the DC-X had completed eight flights in two series, reaching 2500 m. On the eighth flight the aeroshell was cracked in a hard landing.

It was decided to return the vehicle to the Huntington Beach factory for repair and modification to the Delta Clipper Experimental Advanced (DC-XA) configuration. The DC-XA was operated by NASA and the Department of Defense under the Reusable Launch Vehicle program. This had a lightweight graphite-epoxy liquid hydrogen tank and an advanced graphite/aluminum honeycomb intertank built by McDonnell Douglas; an aluminum-lithium liquid oxygen tank built by Energia (Russia); and an improved reaction control system from Aerojet. These improvements reduced dry vehicle mass by 620 kilograms. The flight vehicle, bizarrely renamed by NASA 'Clipper Graham', was tested at White Sands during the summer of 1996, and demonstrated a 26-hour turnaround between its second and third flights, a first for any rocket. After the fourth flight, however, the DC-XA suffered severe damage and the program ended due to lack of funding.

Program Manager for the original SSRT Program was Major Jess Sponable, U.S. Air Force. The prime contractor was McDonnell Douglas Aerospace, Huntington Beach, Calif. The Delta Clipper industrial team consisted of:

• McDonnell Douglas Space Systems Company
• Douglas Aircraft Co., Long Beach, California
• McDonnell Aircraft Co, St. Louis, Mo.
• McDonnell Douglas Electronic Systems Co.
• McDonnell Douglas Missile Systems Co.
• McDonnell Douglas Research Laboratories, St. Louis, Mo.
• Pratt and Whitney Government Engines and Space Division, West Palm Beach, Fla.
• Scaled Composites, Mojave, Calif.
• Aerojet Propulsion Division, Sacramento, California
• Allied Signal Aerospace Co., Torrance, Calif.
• General Connector, San Fernando, California
• Eagle Engineering
• Harris Corp., Rockledge, Fla.
• Honeywell Clearwater, Fla.
• Martin Marietta, Denver, Colo.
• Messerschmitt-Bolkow-Blohm / Deutsche Aerospace, Munich, Germany
• Fluor Daniel
• Process Fabrication, Inc., Santa Fe Springs, Calif.
• Integrated Systems, Santa Clara, Calif.
• Chicago Bridge and Iron Services, Inc., Oak Brook, Ill.
• SpaceGuild, San Carlos, Calif.

• 12 m high, 4.1 m diameter at base, conical shape
• Empty mass: 9100 kg. Fuelled mass: With full load of propellants:18,900 kg
• Propellants: Liquid oxygen and liquid hydrogen
• Propulsion: Four RL-l0A5 rocket engines, each generating 6,100 kgf thrust. Each engine throttleable from 30% to 100%. Each gimbals +/-8 degrees.
• Reaction Controls: Four 440-lb thrust gaseous oxygen, gaseous hydrogen thrusters
• Guidance, Navigation and Control Avionics: Advanced 32 bit, 4.5 mips computer, F-15 Navigation System with ring laser gyros. F/A-18 accelerometer and rate gyro package. Global Positioning Satellite P(Y) code receiver. Digital data telemetry system. Radar altimeter.
• Hydraulic System: Standard hydraulic aircraft-type system to drive vehicle's five aerodynamic flaps and eight engine gimbal actuators (two per engine).
• Construction Materials: Aeroshell and base heat shield: Graphic Epoxy composite with special silicone-based thermal protection coating; Main propellant tanks: 2219 alloy aluminum; Main structural supports: aluminum; Landing gear: steel and titanium

The Chief Technical Advisor for the DC-X has noted that reusable launchers, by their nature have intact abort and that do not crash - they have mishaps not failures. Even the several fires and one auto-abort were not flight test failures. In fact, the auto-abort mishap was the first time that a rocket powered unmanned vehicle had ever saved itself from a fiery death. Even the final test which destroyed the vehicle, met all test objectives and had achieved full engine shut-down prior to the one landing gear collapse (due to a very tired, heat exhausted human technician error).

Development Cost $: 58.900 million in 1991 dollars in 1991 dollars. Total Development Built: 1.

Stage Data - DC-X

• Stage 1. 1 x DC-X. Gross Mass: 16,320 kg (35,970 lb). Empty Mass: 7,200 kg (15,800 lb). Thrust (vac): 262.800 kN (59,080 lbf). Isp: 373 sec. Burn time: 127 sec. Isp(sl): 316 sec. Diameter: 3.05 m (10.00 ft). Span: 3.66 m (12.00 ft). Length: 11.89 m (39.00 ft). Propellants: Lox/LH2. No Engines: 4. Engine: RL-10A-5. Status: Out of Production.

Photo Gallery

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DC-X Credit: NASA

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• Nation: USA. Agency: MDAC. Apogee: 0 km (0 mi). Verified flight control systems and vertical landing capabilities. Reached 46 m altitude in a 59 second flight..

• Nation: USA. Agency: MDAC. Apogee: 0 km (0 mi). Ascent and landing mode control and ground effects survey. Reached 92 m in a 66 second flight..

• Nation: USA. Agency: MDAC. Apogee: 0 km (0 mi). Demonstrated 180 degree roll; provided aerostability data. Reached 370 m during 57 second flight. Last test of first series -- vehicle mothballed when SDIO funding ran out..

• Nation: USA. Agency: MDAC. Apogee: 1.00 km (0.60 mi). First flight of second series after additional SDIO funding received. Full propellent load; radar altimeter in control loop. Reached altitude of 870 m during a 136 second flight..

• Nation: USA. Agency: MDAC. Apogee: 0 km (0 mi). In-flight abort after gaseous hydrogen explosion; vehicle demonstrated autoland capabilities. Reached altitude of 790 m during a 78 second flight..

• Nation: USA. Agency: BMDO. Apogee: 1.00 km (0.60 mi). Continued expansion of flight envelope; constant angle of attack. Reached altitude of 1330 m during a 124 second flight..

• Nation: USA. Agency: BMDO. Apogee: 1.00 km (0.60 mi). First use of reaction control system thrusters; demonstrated change of angle of attack from 0 to 70 degrees. Reached altitude of 1740 m during a 132 second flight..

• Nation: USA. Agency: BMDO. Apogee: 2.00 km (1.20 mi). Final flight of second series. Demonstrated turnaround maneuver. Aeroshell cracked during 4 m/s hard landing. Reached altitude of 2500 m during a 124 second flight..

• DC-XA Rollout - . Nation: USA.

• DC-XA arrives at White Sands - . Nation: USA.

• DC-XA Engine Test 1 - . Nation: USA.

• DC-XA Engine Test 2 - . Nation: USA.