Encyclopedia Astronautica
M-1



satupeng.gif
Advanced Engines
Advanced Engines planned for uprated Saturn and Nova boosters
Credit: © Mark Wade
m1engine.jpg
M1 Engine
Credit: NASA
Aerojet lox/lh2 rocket engine. 5335.9 kN. Study 1961. Isp=428s. Engine developed 1962-1966 for Uprated Saturn and Nova million-pound payload boosters to support manned Mars missions. Reached component test stage before cancellation.

Aerojet, the perennial loser in NASA large engine competitions, finally felt itself a winner when it received the hotly contested contract to develop the M-1 liquid oxygen/hydrogen engine for post-Saturn launch vehicles (Uprated Saturn and Nova). These would be used in the million-pound payload boosters required in the 1970's to support manned Mars missions and lunar colonies. The $238 million contract for the M-l (NASA/LeRC number NAS 3-2555) was issued on April 30, 1962.

The technological leap involved was tremendous - at that time the only Lox/LH2 engine in the flight-test stage was the 6,800 kgf Pratt and Whitney RL10. But Aerojet brought substantial experience to the contract. They had demonstrated a 1,360 kgf Lox/LH2 engine as early as 1948. Starting in 1958 Aerojet had developed a Lox/LH2 version of the Titan 1 LR87 engine with a thrust of 68,000 kgf, although this used a completely different technology from that planned for the M-1. Other related Aerojet work included the 23,000 kgf NERVA nuclear thermal engine. Work had been underway on this since 1960, but it used only liquid hydrogen and did not address combustion or oxygen pumping.

Dan Price directed this remarkable program, which included Dr. John Moise, Hal Campen, and Ken Unmack in key roles. Unmack remembered, ‘The M-l size took months to get used to. We couldn't draw anything full-scale. We were just getting used to all this when the program started winding down'.

Work was slowed down by reduced funding levels within a year of the initial award, due to priority of the basic Saturn V for the Apollo moon landing. Within three years NASA's hopes for a massive post-Apollo program had faded, and the M-1 contract was concluded on August 24, 1965. The last M-1 test was run in August 1966, with cleanup activities extending into 1968. The innovations and technology of the M-1 were helpful to NERVA, and 25 years later, to the STEP shuttle engine improvement programme.

The M-1 was a single chamber LOX/LH, unit designed initially for 545,000 kgf vacuum thrust, and later uprated to 680,000 kgf. It used a simple gas generator cycle with separate hydrogen and oxygen turbopump assemblies. The nozzle expansion ratio of 40:1 was appropriate for altitude operations, and its relatively high chamber pressure of 1000 psi contributed to its vacuum specific impulse of 430 sec. Nominal duration was 400 seconds (with total operational qualified life of 3000 seconds), the unit weighed 9,000 kg and measured 8.15 m in height and 5.39m in diameter. A unique feature was that the nozzle exit cone was cooled by the turbine exhaust gases from the 14:1 area ratio point to the exit. The remainder of the nozzle and the thrust chamber were regeneratively cooled with the liquid hydrogen. Total turbopump shaft power was on the order of 100,000 horsepower. Although the M-1 never reached the phase of all-up engine test, it remains the largest hydrogen/oxygen thrust chamber assembly ever tested. Propellant consumption was over 100 tonnes (two tank car loads) per minute. During hydrogen pump testing it was totally impractical to recapture the pumped hydrogen. Therefore it was simply flared. The first test was at night produced what was the world's largest and most startling flare.

The Lox/LH2 LR87 had used impinging stream injectors and single stage pumps. The M-l had coaxial injectors and over ten stages in the fuel pump.

The hydrogen pump was rated at 75,000 horsepower and was operated at more than 90,000 HP during testing.

The M-1 was a modular engine, and the development proceeded simultaneously on each module. The most significant hardware built and tested included:

  • Thrust chambers (2 uncooled, 6 cooled)
  • Liquid hydrogen turbopump assembly (75,000 horsepower; 4 were in process)
  • Liquid oxygen turbopump (27,000 horsepower - 4 built)
  • Main hydrogen valve, Gas generator (11 units completed and several tested)
  • Injectors (several configurations fabricated, but only one version was tested).

Significant program results included:

  • The feasibility of all major M-l engine components was demonstrated (except for the cooled chamber and the gas cooled skirt).
  • Performance data were obtained and mechanical integrity was established for the injector, fuel turbopump, oxidiser turbopump, and the gas generator assembly
  • The thrust chamber injector demonstrated stable operation and no damage at performance levels in excess of the PFRT requirement
  • Major facilities for testing the fuel and oxidiser turbopumps, as well as the thrust chamber assembly were constructed and successfully operated.

The figures indicated in the specification are for the upper-stage version, which was the contract goal. In first stage applications the engine characteristics would have been:

Vacuum thrust: 544118 kgf / 5335.9 kN
Sea level thrust: 394104 kgf / 3864.8 kN
Mass: 9068 kg
Diameter: 4.28 m
Lenght: 7.72 m
Area ratio: 30
Chamber pressure: 28 atm
Vacuum specific impulse: 428 sec
Sea level specific impulse: 310 sec

Thrust (sl): 3,864.800 kN (868,842 lbf). Thrust (sl): 394,104 kgf. Engine: 9,068 kg (19,991 lb). Chamber Pressure: 69.00 bar. Area Ratio: 28. Thrust to Weight Ratio: 60.

Status: Study 1961.
Unfuelled mass: 9,068 kg (19,991 lb).
Height: 7.72 m (25.32 ft).
Diameter: 4.28 m (14.04 ft).
Thrust: 5,335.90 kN (1,199,558 lbf).
Specific impulse: 428 s.
Specific impulse sea level: 310 s.
Burn time: 250 s.

More... - Chronology...


Associated Countries
See also
Associated Launch Vehicles
  • Nova 4S American heavy-lift orbital launch vehicle. NASA Nova design using a cluster of 4 x 240 inch solid motors used as first stage; upper stages as Nova 7S and 8L. More...
  • Nova 7S American heavy-lift orbital launch vehicle. NASA Nova design using a cluster of 7 x 160 inch solid motors used as first stage; upper stages as Nova 4S and 8L. More...
  • Nova 8L American heavy-lift orbital launch vehicle. Most capable NASA Nova design, studied in June 1960 just prior to selection of Saturn for moon landing. Used a three stage configuration of eight F-1 engines in stage 1, two M-1 engines in stage 2, and one J-2 engine in stage 3. Similar to the Saturn C-8 except in the use of M-1 engines. Unlike other modular Nova designs of the time, this one had the unitary stage construction of Saturn. More...
  • Nova MM 14A American heavy-lift orbital launch vehicle. Nova design using 4 300 inch solids as first stage, 5 M-1 in second stage. Operational date would have been April 1973 More...
  • Nova MM 1B American heavy-lift orbital launch vehicle. Nova design using existing engines; 14 F-1A in the first stage, 2 M-1 in the second. Operational date would have been December 1972 More...
  • Nova MM 14B American heavy-lift orbital launch vehicle. Nova design using 4 280 inch solids as first stage, 4 M-1 in second stage. Operational date would have been February 1973 More...
  • Nova MM 1C American heavy-lift orbital launch vehicle. Nova design using existing engines; 18 F-1A in the first stage, 3 M-1 in the second. Operational date would have been February 1973 More...
  • Nova-1 DAC American heavy-lift orbital launch vehicle. Douglas/Bono design for Nova using Lox/RP-1 in first stage, existing engines. More...
  • Nova-2 DAC American heavy-lift orbital launch vehicle. Douglas/Bono design for Nova using LH2/Lox in both stages. More...
  • Nova GD-E American heavy-lift orbital launch vehicle. General Dynamics Nova design using 325 inch solid motors as first stage, M-1 engines in second stage. Recoverable solid motors, separation at 1,972 m/s at 53,000 m altitude; splashdown using retrorockets under 3 61 m diameter parachutes 610 km downrange. Recovery of solid motors forshadowed same approach on shuttle 15 years later. Masses estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova GD-F American heavy-lift orbital launch vehicle. General Dynamics Nova design using new 3.5 million kgf Lox/Kerosene engines in first stage. Recoverable stage; separation at 3,365 m/s at 89,300 m altitude; splashdown using retrorockets under 8 46 m diameter parachutes 1300 km downrange. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova GD-J American heavy-lift orbital launch vehicle. General Dynamics Nova design using recoverable Lox/RP-1 stage of ballistic shape with 3 million kgf engines; separation at 3,420 m/s at 93,900 m altitude; splashdown using retrorockets under 7 parachutes 1340 km downrange. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...

Associated Manufacturers and Agencies
  • Aerojet American manufacturer of rockets, spacecraft, and rocket engines. Aerojet, Sacramento, CA, USA. More...

Associated Propellants
  • Lox/LH2 Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today. More...

Bibliography
  • Dorman, Bernie, et. al., Aerojet: The Creative Company, Stuart F Cooper Company, Los Angeles, 1995..

Associated Stages
  • Astro-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 302,183/32,558 kg. Thrust 8,820.00 kN. Vacuum specific impulse 410 seconds. Engines 1 x M-1 plus 2 x J-2 More...
  • Nova MM 1B-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,361,000/122,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been February 1973 More...
  • Nova MM 14B-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,721,000/245,000 kg. Thrust 26,683.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been February 1973 More...
  • Nova MM 14A-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 3,401,000/295,000 kg. Thrust 33,352.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been April 1973 More...
  • Nova GD-J-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,270,000/91,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova GD-F-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,497,000/91,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova GD-E-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 4,535,000/363,000 kg. Thrust 26,683.00 kN. Vacuum specific impulse 428 seconds. Massed estimated based on tank volumes, total thrust, and first stage burnout conditions. More...
  • Nova DAC-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,268,000/143,000 kg. Thrust 26,683.00 kN. Vacuum specific impulse 426 seconds. Operational date would have been July 1977. Recoverable stage. More...
  • Nova DAC 2-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 1,633,000/113,000 kg. Thrust 13,346.00 kN. Vacuum specific impulse 426 seconds. Operational date would have been July 1977. Recoverable stage. More...
  • Nova 60-8-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 680,000/54,000 kg. Thrust 10,669.00 kN. Vacuum specific impulse 428 seconds. Mass estimated based on total LV weight. J-2-powered version of this stage also proposed. More...
  • Nova MM 1C-2 Lox/LH2 propellant rocket stage. Loaded/empty mass 2,041,000/163,000 kg. Thrust 20,015.00 kN. Vacuum specific impulse 428 seconds. Operational date would have been February 1973 More...
  • Nova DAC 2-1 Lox/LH2 propellant rocket stage. Loaded/empty mass 4,943,000/311,000 kg. Thrust 102,293.00 kN. Vacuum specific impulse 410 seconds. Operational date would have been July 1977. Recoverable stage. 10% plug nozzle. More...
  • SLS Stage C Lox/LH2 propellant rocket stage. Loaded/empty mass 825,000/58,000 kg. Thrust 10,673.32 kN. Vacuum specific impulse 428 seconds. Launch vehicle core stage for Project Lunex. Masses estimated based on optimum apportioning of B+C stage total masses. Thrust, engines estimated based on requirements. More...

Home - Browse - Contact
© / Conditions for Use