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Houston, Texas February 7, 1966 Maprogram Ap0l);O Xorking Pax3 No Copy No. ti NiUA PROGRAM APSLLO WOP?GPAPER NO. 1195 &L??OLLO SPACECRAFT LlQuzD PRIMARY PROPMON HYSTEMS NASA \' NATIGNAL AERONAUTICS AND SPACE ADMINISTRATION MANNED SPACECRAFT CENTER HOUSTON, TEXAS FEBRUARY 7, 1966 MAPROGRAM AP0L);O XORKING PAX3 NO. 11% APOLLO SPACECRAFT LIQUID PRlMARY PROPULSION SYSTEMS PEU3PARED BY Propulsion Analysis Section Propulsion Analysis Section M G?L. Spencer Propulsion Analysis Section PrEmsion AnaJysis Section J. C. Smithson Propulsion Analysis Section AUTHORIZED FOR DISTRIBUTION L Assistant Director for Engineering and Developent NATIONAL AERONAUTICS AND SPACE A.IMINISTRATION MANNED SPACECRAFT CENTER HOUSTON, TEXAS FEBRUARY 7, 1966 iii PRECEDING PAGE BLANK NOT CONTENTS Section Page S~............................. 1 Mission Requirements . 1 Design ........................... 2 Philosophy 0.. 0, * . 0 0 0 0 0 0 0 . 0 2 Basic features . 2 Pressurization and Propellant Feed Systems . Chamber-Injector Compatibility and Ablation . SrnVICE PR0PuLsI0N SYSTEM. 0 0 0 0 0 0 . 0 . 0 0 . 0 Operation. ......................... 4 DESCENT AND ASCENT PR0puI;sION SYSTEMS . Descent Propulsion System Operation , . Ascent Propulsion System Operation . , . TAuLE I - APOLLO SPACECRAFT ENGIW SPECIFICATIONS 0 . 0 . 0 0 iv FIGURES Figure 1 Earth vicinity and translunar mission ......... 10 2 Lunar vicinity mission ................ ll 3 Transeazth and entry mission ............. 12 4 The Apollo spacecraft ................. 13 5 Service propulsion system ............... 14 6 SPEengine ...................... 15 7 Lunar excursion module ................ 16 8 Descent and ascent engines as installed in LEN . 17 9 Descentstage ..................... 18 10 Ascentstage ..................... 19 u. LEMdescentengine .................. 20 12 LEMascentengine ................... 21 13 SPS thrust chamber .................. 22 14 DPS thrust chamber .................. 23 15 Ascent engine chamber detail ............. 24 16 SPS nozzle extension ................. 25 17 Descent engine cros6 section ............. 26 18 Ascent engine cross section .............. 27 19 SPSinjector ..................... 28 20 Ascent engine injector ................ 29 21 Descent engine injector detail ............ 30 Figure Page 22 SPS pressurization and propellant system . 31 23 SPS propellant valve assembly . 32 24 SPS propellant flow . 33 25 Descent stage supercritical helium system . 3 26 Descent engine flow control . 35 27 Descent engine flow control system . , . 36 28 Descent engine flow control valve . 37 29 Descent engine thro-btle concept . 38 30 Ascent engine pressurization and propellant system . 39 31 Ascent engine flow control . 40 APOLLO SPACECRAFT LIQUID PRIMARY PROPULSION SYSTEMS By Propulsion Analysis Section The purpose of this document is to provide a general faniliariza- tion with the Apollo spacecraft liquid primary propulsion systems. A brief discussion of the mission, '."rgine history, and design philosophy is presented for the three primary ApoUo spacecraft engines, including the specifications and gecunetry ol" each. Mission Requirements The Pgollo mission will place three men into lunar orbit, land two of these men on the lunar surface, and return the three men to earth. The entire mission will take approximately 8 days and is shorn in detail in figures 1 through 3. The p pol lo spacecraft (fig. 4), consisting of the service module, the caranand module, and the two-stage lunar excursion module, contains three primary propulsion systems, which are the subject of this report. The service module contains the service propulsion systen, (SPS) (fig. 5). The service module engine (fig. 6) has a naminal thrust of 21 !jOO pounds and.is us& for midcourse correction during the translunar phase, lunar ofiit insertion, transearth injection, and midcourse cor- rection during the transearth phase. The lunar excursion module (W)(figs. 7 through 10) contains the descent and ascent propulsion systems (DFQ and APS). The LEI4 descent engine (fig. U)has a variable thrust capability between 10 500 and 1 050 pounds and provides the tbrust for retrofire fran lunar orbit and lunar descent, hover, and landing. The LEM ascent engine (fig. 12) ha8 a naninal thrust of 3500 pound8 and provldes the thrust to place the ascent stage of the LEM into lunar orbit for rendezvous With the orbiting coIIpD.md and service lllOdUle. 2 Design Fbi1osoyhy.- One of the predaninant requirements which dictated the designs of the Apollo engines is crew safety, and, therefore, propulsion systea, reliability. To achieve the desired system reliability, the fol- lowing design features were incorporated in all three Apollo primary pro- pulsion systems. 1. The propellants are storable at ambient temperature, eliminating the problems associated with long term storage and venting of cryogenic propellants in space under zero gravity conditions. 2. The combustion chambers are ablatively cooled because of the rugged nature of ablatives and their high resistance to sudden failure. 3. The engines are pressure-fed to eliminate complexities asso- ciated with pump fed engines. 4. Redundant cmponents are used for all rn9~ngparts in the enghe pressurization and propellant feed systems. This provides for mission campletion or a safe mission abort if a malfbction occurs in one leg of a redundant component or in a number of redundant components in the pro- pulsion system. Basic features.- The design philosophy and mission requiremeats of Apollo have dictated a number of design features which are common to the three primary spacecraft propulsion systems. All three primary propul- sion systems (figs. 13, 14 and 15) use ablative thrust chambers, have re- dulldant valves, and are pressure fed. Optimum performance and minimum weight are basic requirements because of the nature of the Apllo mis- sion. Large expansion ratio nozzles are used for optimum performance in the vacuum conditions of space. Relatively low chamber pressure and low pressure drops through the propellmt feed system and injector are re- quired for optimum vehicle efficiency with the pressurized propellant feed system. The propellants utilized in these systems are earth stor- able and hypergolic. The fuel is a blended hydrazine (approximately 50 percent unsymmetrical dimethywazine and 50 percent anhydrous hy- . drazine). The oxidizer is nitrogen tetroxide. These propellants provide the performance necessary to accomplish the Apollo mission and have ac- ceptable handling and material compatibility characteristics. To minimize engine weight, the service module and descent engines have major portions of their nozzles constructed of high ternperatwe Eetals (figs. 16 and 17) which are cooled by radiant heat transfer to space, Tke entire nozzle of the ascent engine (fig. 18) is constructed 3 of ablative material because the required engine mounting location and vehicle heat shield configuratioa does not provide a sufficient view factor for radiation to space. The LE24 descent engine utilizes a variable area injector and vwi= able area cavitating venturi flow control valves for throttling. ‘fie single-injector and single-fuel and oxidizer-flow control valves aze ex= ceptions to the rule of redundancy for all moving comFnents. However, the actuation system which controls these componats is redundant . Pressurization and Propellant Feed Systems Gaseous helium is used for pressurization, with the gas being stored at ambient temperature in the SPS and APS, and at cryogenic temperatures in the DPS. The DPS cryogenic helium storage system takes advantage of the weight savings possible in the helium storage vessel due to increased helium density at reduced temperatures, but results in increased complex- ity in the pressurization system. In each propulsion system, oxidizer and fuel tasks are fed fmn a common pressure regulation assdly to guarantee equal ullage pressures in the oxidizer and fie1 tanks. A series-parallel valve arrangement is used on the pressure regula- tion, check valve, and shutoff valve assemblies of each propulsion sys- tem. In the evcnt a valve should fail open, the series arrangement provides a second valve to shut off flow. Ia the event a valve should fail closed, the parallel arrangement provides a second flow path for continued operation. The oxidizer and fuel engine valves are paired so that each set is mechanically linked and operates fram a common actuator assembly to guarantee the proper relative valve timing during operation. Trim orifices are.located in the propellant feed systems to trim the pro- pulsion system and to provide the desired mixture =%io and thrust level. Chamber-Injector Compatibility and Ablation b%amber-injector compatibility is a mador concern when optimum en- girle efficiency is a design requirement. High performance requires high combustion temperatures which can result in high ablation rates. High ablation results in a reduced nozzle =ea ratio, a performance reduction, and the requirement for a thicker and heavier ablative chamber. To con- trol ablation, a barrier of lcw mixture ratio cambustion gases is used to isolate the hot core of cambustion gas from the chamber wall. The ser- vice module engine Sdector (fig. 19) uses a row of showerhead fuel ori- scent to the chamber wall, while the IZM ascent engine injector uses a low mixture ratio unlike-doublet Wacent to the chamber 4 wall. Because of the unique design of the variable area descent engine injector (fig. 21), no specific portion of the fuel can be considered as the film coolant. SERVICE ~OPULSIONSYS!lXb! Operation The helium pressurization supply
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