Aerojet M-1
This article includes a list of general references, but it lacks sufficient corresponding inline citations. (June 2012) |
Gas-generator cycle | |
Performance | |
---|---|
Thrust, vacuum | 1,500,000 lbf (6.67 MN) |
Thrust-to-weight ratio | 60 |
Chamber pressure | 1,000 psia |
Specific impulse, vacuum | 428 seconds (4.20 km/s) |
Dimensions | |
Length | 7.72 m |
Diameter | 4.28 m |
Dry weight | 9068 kg |
The
History
The M-1 traces its history to US Air Force studies from the late 1950s for its launch needs in the 1960s. By 1961 these had evolved into the
The smallest model, intended to launch the
The SLS also envisioned a number of much larger designs intended to launch the Air Force's
When
However, the Apollo spacecraft requirements quickly grew, settling on a 10,000 lb (4,500 kg) spacecraft (the
In 1961, President
Uprating thrust, then cancellation
With the selection of Saturn for the lunar missions, work on Nova turned to the post-Apollo era. The designs were re-targeted for crewed planetary expeditions, namely a crewed landing on Mars. Even utilizing a lightweight mission profile like that selected for Apollo, a Mars mission required a truly massive payload of about one million pounds to low Earth orbit. This led to a second series of design studies, also known as Nova, although they were essentially unrelated to the earlier designs.
Many of the new designs used the M-1 as their second-stage engine, although demanding much higher payloads. In order to meet these goals, the M-1 project was uprated[when?] from 1.2 million pounds force to a nominal 1.5 million pounds force, and the designers deliberately added more turbopump capability to allow it to expand to at least 1.8 million and potentially up to 2.0 million pounds force.[1] Additionally, the M-1 was even considered for a number of first-stage designs, in place of the F-1 or the 180-inch (4,600 mm) solids. For this role the specific impulse was dramatically reduced, and it appears that some consideration was given to various expanding nozzle designs to address this.
M-1 development continued through this period, although as the Apollo program expanded, NASA started cutting funding to the M-1 project in order to complete Saturn-related developments first. In 1965, another NASA project[
By 1966 it was clear that present funding levels for NASA would not be maintained in the post-Apollo era. The Nova design studies ended that year, and the M-1 along with it. The last M-1 contract expired on August 24, 1965, although testing continued on existing funds until August 1966. Studies on the J-2T ended at the same time. Although the HG-3 was never built, its design formed the basis for the
The final report (1966)[2] found:
- The feasibility of all major M-1 Engine components, except for the cooled chamber and the gas-cooled skirt, was demonstrated.
- Performance data were obtained and the mechanical integrity was established for the injector, the fuel turbopump, the oxidizer turbopump, and the gas generator assembly. Also, it was established that these components are satisfactory for use in a demonstration engine.
Prototypes
Over the three-year lifetime of the project, a total of eight combustion chambers were built (two of them uncooled test units), eleven gas generators, four oxygen pumps, as well as four hydrogen pumps that were in the process of being completed.
Scaled down models of the pumps were used during design/development to 1963.[3]
Description
The M-1 used the
In most American designs, a gas-generator engine would dump the exhaust from the turbines overboard. In the case of the M-1, the resulting exhaust was relatively cool, and was instead directed into cooling pipes on the lower portion of the engine skirt. This meant that liquid hydrogen was needed for cooling only on the high-heat areas of the engine—the combustion chamber, nozzle and upper part of the skirt—reducing plumbing complexity considerably. The gas entered the skirt area at about 700 °F (371 °C), heating to about 1,000 °F (538 °C) before being dumped through a series of small nozzles at the end of the skirt. The exhaust added 28,000 lbf (120 kN) of thrust.
The engine was started by rotating the pumps to operating speed using helium gas stored in a separate high-pressure container. This started the fuel flow into the main engine and gas generator. The main engine was ignited by a spray of sparks directed into the combustion chamber from a pyrotechnic device. Shutdown was achieved by simply turning off the fuel flow to the gas generator, allowing the pumps to slow down on their own.
The use of separate turbopumps and other components allowed the various parts of the M-1 to be built and tested individually.
Combustion chamber and injectors
- Thrust: 1.5 M lb (at 200,000 ft)[4]
- Thrust chamber pressure: 1,000 psia,[1]: Table 1 1,200 psia for 1.8 M lb version
- Thrust chamber diameter: 42 inches
- Thrust chamber material: 200 tubes of 347 stainless steel, in an Inconel 718 bolt-on jacket.[5]
- Injector type: coaxial
- Injector body material: 347 stainless steel[5]
- Number of Injector elements: 1,200 to 3,000 anticipated[5]
- Nozzle throat diameter:
Gas generator
- Burns 110 lb/s (oxidiser:fuel, 0.8)
- Exhaust pressure: 1100 psi
- Exhaust temperature: 1,000 °F (538 °C)
- Gas generator exhaust fed back into lower nozzle for cooling
LOX turbopump
- Axial flow
- RPM: 36,700[6]
- Input pressure: 30 ft[6] (of LOX)
- Pressure increase: 3,400 ft[6] (of LOX); i.e., 1,700 psi[7]
- Flow rate: up to 3,000 lb/sec, 2,921 lb/s nominal [4]
- Axial thrust load on bearings: in excess of 30,000 lb[3]
- Bearings: oxygen lubricated, 440C stainless steel balls, with "glass filled Teflon cages"[3]
LH2 turbopump
- 2 stage turbine with 10 stage axial flow pump[8]
See also
References
- ^ a b Dankhoff 1963, pp. 1–2.
- ^ Development of a 1,500,000-lb-thrust /Nominal Vacuum/Liquid Hydrogen/Liquid Oxygen Engine Final Report, 30 Apr. 1962 - 4 Aug. 1966
- ^ a b c Dankhoff 1963, p. 11.
- ^ a b Dankhoff 1963, p. 6.
- ^ a b c Dankhoff 1963, p. 8.
- ^ a b c Dankhoff 1963, p. 10.
- ^ Dankhoff 1963, p. 25.
- ^ Dankhoff 1963, p. 9.
Bibliography
- Dankhoff, Walter F. (October 1963). The M-1 Rocket Engine Project (PDF). Washington, D.C.: NASA. Archived from the original (PDF) on 2015-01-05.
- Mechanical design of the M-1 axial flow liquid hydrogen fuel pump
- Development of a 1,500,000-lb-thrust /nominal vacuum/ liquid hydrogen/liquid oxygen engine Final report, 30 Apr. 1962 - 4 Aug. 1966/ NASA document covering the M-1 Project from inception to completion. 406p
- Activation and initial test operations, large rocket engine - Turbopump test facilities Technology report Aerojet General report on the development of the test facilities for the M-1 Turbopump
- Activation and Initial Test Operations, Large Rocket Engine - Thrust Chamber Test Facilities Technology Report Aerojet General report on the development of the test facilities for the M-1 Thrust Chamber
- Development of LO2/LH2 Gas Generators for the M-1 Engine NASA document covering the development of the gas generators for the M-1 engine
- Development of liquid oxygen/liquid hydrogen thrust chamber for the M-1 engine NASA document covering the development of the M-1 thrust chamber
- Design study of modification of m-1 liquid hydrogen turbopump for use in nuclear reactor test facility
- Analytical and experimental vibration analysis of the turbine buckets for the M-1 liquid oxygen turbopump
- Economic analysis of Perlite versus super insulation in liquid hydrogen storage and run vessels for the M-1 Program
- Aerodynamic design and estimated performance of a two-stage Curtis turbine for the liquid oxygen turbopump of the M-1 engine
- Investigation of the starting characteristics of the M-1 rocket engine using the analog computer
- Analysis of the M-1 liquid hydrogen turbopump shaft critical whirling speed and bearing loads
- Cold-air performance evaluation of scale model oxidizer pump-drive turbine for the M-1 hydrogen-oxygen rocket engine. I - Inlet feedpipe-manifold assembly
- Cold-air performance evaluation of scale model oxidizer pump-drive turbine for the M-1 hydrogen-oxygen rocket engine. II - Overall two-stage performance
- Cold-air performance evaluation of scale model oxidizer pump-drive turbine for the M-1 hydrogen-oxygen rocket engine. III - Performance of first stage with inlet-feedpipe-manifold assembly
- Cold-air performance evaluation of scale model oxidizer pump-drive turbine for the M-1 hydrogen-oxygen rocket engine. IV - Performance of first stage with modified inlet feedpipe-manifold assembly
- Design and development of liquid hydrogen cooled 120mm roller, 110mm roller, and 110mm tandem ball bearings for M-1 fuel turbopump
- Valve lipseals M-1 sleeve-type thrust chamber valve
- Development of liquid oxygen cooled 110MM roller and tandem ball bearings at up to .5 x 106 DN values for the oxidizer turbopump of the M-1 engine Technology report
- Aerodynamic design - Model II turbine M-1 fuel turbopump assembly
- Analysis and experimental verification of axial thrust on the M-1 liquid oxygen turbopump
- M-1 engine test complex data acquisition systems
- The mechanical design of a two-stage impulse turbine for the liquid hydrogen turbopump of the M-1 engine
- Summary of observed results when chilling the M-1 fuel turbopump to liquid hydrogen temperature
- Mechanical design of a Curtis turbine for the oxidizer turbopump of the M-1 engine
- Hydraulic design of the M-1 liquid hydrogen turbopump
- Summary of materials technology of M-1 engine
- Cooled baffle development for M-1 engine using a subscale rocket engine
- M-1 injector development - Philosophy and implementation
- Cold-air performance evaluation of a scale- model fuel pump turbine for the M-1 hydrogen-oxygen rocket engine
- Application of Alloy 718 in M-1 engine components
- M-1 engine subscale injector tests
- Scale model study of flow patterns in the inlet manifold of the fuel pump drive turbine for the M-1 hydrogen-oxygen rocket engine
- M-1 injector development - Philosophy and implementation
- Hydrogen gas pressure vessel problems in the M-1 facilities
- Spin Test of Turbine Rotor NASA Contractor Report into spin tests of the turbine built for the M-1 Oxidizer Turbopump dated February 1972