Jump to content

Aerojet M-1

From Wikipedia, the free encyclopedia

This is an old revision of this page, as edited by Joaquin89uy (talk | contribs) at 13:39, 14 July 2022 (Adding national origin.). The present address (URL) is a permanent link to this revision, which may differ significantly from the current revision.

M-1
M-1 rocket engine specifications
Country of originUnited States
First flightDevelopment abandoned at pre-prototype stage
ManufacturerAerojet
StatusAbandoned in development stage.
Liquid-fuel engine
PropellantLOX / liquid hydrogen
CycleGas-generator cycle
Performance
Thrust, vacuum1,500,000 lbf (6.67 MN)
Thrust-to-weight ratio60
Chamber pressure1,000 psia
Specific impulse, vacuum428 seconds (4.20 km/s)
Dimensions
Length7.72 m
Diameter4.28 m
Dry mass9068 kg

The Aerojet M-1 was the largest and most powerful liquid-hydrogen-fueled liquid-fuel rocket engine to be designed and component-tested. It was originally developed during the 1950s by the US Air Force. The M-1 offered a baseline thrust of 6.67 MN (1.5 million lbf) and an immediate growth target of 8 MN (1.8 million lbf). If built, the M-1 would have been larger and more efficient than the famed F-1 that powered the first stage of the Saturn V rocket to the Moon.

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 Space Launcher System design. The SLS consisted of a series of four rocket designs, all built around a series of solid-fuel boosters and liquid-hydrogen-powered upper stages.

The smallest model, intended to launch the Dyna-Soar, used two 100-inch (2,500 mm) solids and an "A" liquid core. To power the "A" booster, Aerojet was contracted to convert an LR-87, used in the Titan II missile, to run on liquid hydrogen. A prototype was successfully tested between 1958 and 1960. Initial studies of the 100-inch (2,500 mm) solid were also handed to Aerojet, starting in 1959.

The SLS also envisioned a number of much larger designs intended to launch the Air Force's Lunex Project crewed lunar landing. Lunex was a direct landing mission, in which a single very large spacecraft would fly to the Moon, land, and return. In order to launch such a design to low Earth orbit (LEO), a very large booster with a 125,000 lb (57,000 kg) payload would be required. These larger SLS designs followed the same basic outline as the smaller Dynasoar booster, but used much more powerful 180-inch (4,600 mm) solids and the "B" and "C" liquid stages. To provide the required power, the liquid stages mounted a cluster of twelve J-2s. To reduce this complexity, the Air Force also had Aerojet start studies of a much larger hydrogen-fueled design that would replace the twelve J-2s with only two engines. These initial studies would eventually emerge as the M-1, with a thrust of 1.2 million pounds force.

When NASA formed in 1958, they also started planning for a lunar landing. Like the Air Force, their Project Apollo initially favoured a direct ascent profile, requiring a large booster to launch the spacecraft into LEO. Prior to NASA taking over Wernher von Braun's Saturn work for the US Army, they had no large rocket designs of their own, and started a study program known as Nova to study a range of options. Initially, the payload requirements were fairly limited, and the favoured Nova designs used a first stage with four F-1 engines and a payload of about 50,000 lb (23,000 kg). These designs were presented to President Dwight D. Eisenhower on January 27, 1959.

However, the Apollo spacecraft requirements quickly grew, settling on a 10,000 lb (4,500 kg) spacecraft (the CSM) with a three-person crew. To launch such a craft to the Moon required a massive 125,000 lb (57,000 kg) payload to LEO. Nova designs of this capability were quickly presented with up to eight F-1 engines, along with much more powerful upper stages that demanded the M-1 engine. Thus, for a brief period, the M-1 was used on the baseline designs for both NASA's and the Air Force's lunar programs.

In 1961, President John F. Kennedy announced the goal of landing a person on the Moon before the decade was out. After a brief argument, NASA won the mission over the Air Force. However, Nova would require massive manufacturing capability that did not currently exist, and it was not clear that booster construction could be started in time for a landing before 1970. By 1962 they had decided to use von Braun's Saturn V design, which went through a process of re-design to produce a usable booster that could be built in the existing facilities at Michoud, Louisiana.

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[which?] studied advanced versions of the Saturn, replacing the cluster of five J-2s on the S-II second stage with one M-1, five J-2Ts (an improved version of the J-2 with an aerospike nozzle), or a high-pressure engine known as the HG-3, which would later become the direct predecessor of the Space Shuttle's SSME.

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 Space Shuttle Main Engine.

The final report (1966)[2] found:

  1. The feasibility of all major M-1 Engine components, except for the cooled chamber and the gas-cooled skirt, was demonstrated.
  2. 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 gas-generator cycle, burning some of its liquid hydrogen and oxygen in a small combustor to provide hot gases for running the fuel pumps. In the case of the M-1, the hydrogen and oxygen turbopumps were completely separate, each using their own turbine, rather than running both off a common power shaft. The hydrogen and oxygen pumps were some of the most powerful ever built at the time, producing 75,000 horsepower for the former, and 27,000 hp (20,000 kW) for the latter.

M-1 rocket engine display at Evergreen Aviation & Space Museum
A turbopump designed and built for the M-1 rocket engine

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

  1. ^ a b Dankhoff 1963, pp. 1–2.
  2. ^ Development of a 1,500,000-lb-thrust /Nominal Vacuum/Liquid Hydrogen/Liquid Oxygen Engine Final Report, 30 Apr. 1962 - 4 Aug. 1966
  3. ^ a b c Dankhoff 1963, p. 11.
  4. ^ a b Dankhoff 1963, p. 6.
  5. ^ a b c Dankhoff 1963, p. 8.
  6. ^ a b c Dankhoff 1963, p. 10.
  7. ^ Dankhoff 1963, p. 25.
  8. ^ 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
Retrieved from "https://en.wikipedia.org/enwiki/w/index.php?title=Aerojet_M-1&oldid=1098160136"