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The Eagle has Landed

I gasped when I looked out the window and saw this enormous crate at the front door.


This Apollo era rocket engine was originally designed for the Lunar Module Descent Engine (LMDE or DPS), and then employed as the 2nd-stage engine on the Delta Space Launch Vehicle. This may be the only complete unit in existence.


During the Apollo 13 emergency, the LMDE brought the spacecraft back to earth from the moon in an untested manner. Since an earlier explosion took out the main oxygen tanks, they improvised and used the LMDE— the engine from the lunar lander, designed to slow its decent to the moon — to instead push the crippled Command Module and reentry capsule for the return burn and subsequent course corrections. (Here are some Interesting details from Lovell and Haise)


This engine uses hypergolic fuels – noxious chemicals that combust on contact and make for simple, reliable engines that you can use repeatedly in the vacuum of space. The pump for these fuels on the LM ascent engine was the focal point of Neil Armstrong’s nightmares before the Apollo 11 launch.


I’ll post some unpacking photos and details as I gather them below.


Big thanks to Spaceaholic for the following details, engraved in a plaque that now hangs above it in our lobby:


“TRW TR-201 Bipropellant Rocket Engine. The thrust chamber was initially developed for the Apollo Lunar Module and was subsequently adopted for the Delta Expendable Launch Vehicle 2nd stage. The engine made 10 flights during the Apollo program and 77 during its Delta career between 1974-1988. This TR-201 has been configured as a fixed thrust version of the Lunar Module Descent Engine (LMDE) for Delta's stage 2. Multi start operation is adjustable up to 55.6 kN and propellant throughput up to 7,711 kg; and the engine can be adapted to optional expansion ratio nozzles. Development of the innovative thrust chamber and pintle design is credited to TRW Aerospace Engineer Dr. Peter Staudhammer.


The combustion chamber consists of an ablative-lined titanium alloy case to the 16:1 area ratio. Fabrication of the 6A1-4V alloy titanium case was accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. The ablative liner is fabricated in two segments and installed from either end. The shape of the nozzle extension (not installed on the example in this collection) is such that the ablative liner is retained in the exit cone during transportation, launch and boost. During engine firing, thrust loads force the exit cone liner against the case. The titanium head end assembly which contains the Pintle Injector and propellant valve subcomponents is attached with thirty-six A-286 steel ¼ inch bolts.


In order to keep the maximum operating temperatures of the titanium case in the vicinity of 800 degrees (F), the ablative liner was designed as a composite material providing the maximum heat sink and minimum weight. The selected configuration consisted of a high density, erosion-resistant silica cloth/phenolic material surrounded by a lightweight needle-felted silica mat/phenolic insulation.


The installed Pintle Injector, unique to TRW designed liquid propulsion systems, provides improved reliability and less costly method of fuel oxidizer impingement in the thrust chamber then conventional coaxial distributed-element injectors typically used on liquid bipropellant rocket engines.


Number flown: 77 (Delta 2000 configuration)

Dry mass: 300 pounds (with Columbian Nozzle Extension Installed)

Length: 51 inches - Gimbal attachment to nozzle tip (minus nozzle extension)

Maximum diameter: 34 inches (minus nozzle extension)

Mounting: gimbal attachment above injector

Engine cycle: pressure fed (15.5 atm reservoir)

Fuel: 50/50 N2O4/UDMH at 8.92 kg/s

Oxidizer: Dinitrogen tetroxide at 5.62 kg/s

O/F ratio: 1.60

Thrust: 42.923 kN vac

Specific impulse: 303 s vacuum

Expansion ratio: 16:1, 43:1 (with Expansion Nozzle)

Cooling method: Film cooled (upper thrust chamber); quartz phenolic chamber ablation (lower thrust chamber) and columbium (niobium) nozzle radiation (Nozzle extension)

Chamber pressure: 7.1 atm

Ignition: hypergolic, started by 28 V electrical signal to on/off solenoid valves

Burn time: 500 s for total of 5 starts; 10 350 s single burn”

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Taken on March 24, 2010