Nozzle and chamber, phenolic impregnated silica with fiberglass overwrap; nozzle, sprayed zirconia coating; propellant inlets and valves, metal
Thompson Ramo Wooldridge (TRW) Inc. Rocket Engine developed for application as the Apollo Saturn V Launch Vehicle S-IVB (Third Stage) Auxiliary Propulsion System (APS) Attitude Control Engine. Three 150 pound thrust (670 Newton) TRW, Inc. TR-204 engines were employed in each of the two attitude control modules affixed to the S-IVB stage (employed as both the second stage on the Saturn IB and the third stage of the Saturn V launch systems). Hypergolic bi-propellants (Nitrogen Tetroxide combined with Monomethylhydrazine) fed at an oxidizer to fuel ratio of 1.60 was supplied to the engine under positive pressurization via independant helium and propellant storage tanks housed within each APS module. The APS provided roll control during J-2 engine powered phases of flight, pitch, yaw, and roll (attitude) control during coast periods; maneuvering impulse during undocking and extraction of the Lunar Module, and also propellant settling (ullage). Firing control for the TR-204 was initiated by the Instrument Unit Flight Control Computer. As a final act of glory, the APS was continuously fired until depletion to propel the spent S-IVB stage to lunar impact to produce seismic energy for detection by sensors emplaced on the moon's surface during preceding Apollo missions.
The TR-204 combustion chamber is lined with ablatively refrasil material. The engines have quadruple propellant injector valves for redundancy. The thrust chamber assembly (TCA) is integrally fabricated and composed of three major elements: the combustion chamber, the nozzle throat section and the nozzle expansion cone with an expansion ration of 33.9 to 1. The engine was qualified for a total pulse operation of 300 seconds. During the Saturn program, this engine achieved an overall reliability design goal of 0.992 at a 90-percent confidence level while operating.
The engine propellant valve assembly consists of eight normally closed, quad redundant propellant valves (four oxidizer and four fuel), arranged in two series parallel arrangements. Dual failure within the manifold fuel or oxidizer arrangement was required to cause failure of the rocket engine assembly. An assembly closed failure would have prevented any operation while an assembly open failure would have resulted in continuous flow and loss of all propellant. Assembly valve failure cannot occur unless two valves fail open in series or two valved failed closed in parallel. This high degree of reliability was nessessary to prevent the possibility of a mission abort do to loss of spacecraft/launch vehicle attitude control prior to separation of the Saturn third stage and the Apollo Command Service Module (CSM). The inlet ot the valve package contains a 100 mesh screen (150 microns) to protec the engine from large particulate conamination.
The injector valves provided positive on-off control of propellant flow upon command from an external power source (28 volts DC). Four valves, integral in an assembly were capable of simultaneous operation and were synchronized to allow or terminate propellant flow within 3 milliseconds of each other. The opening time for each valve assembly, defined as the time from initiation of open signal to fully open valve package did not exceed 23 milliseconds.
The injector consists of twelve pairs of unlike-on-unlike doublets arranged to minimize hot spots in the combustion chamber. The valve side of the injector is filled with a silver braze heat sink which reduces injector operating temperature. During the 300-second life requirement, the external wall temperature was designed not to exceed 600 degrees F and the maximum valve body external temperature did not exceed 165 degrees F.