Apollo Lunar Module Propulsion System

The Main Propulsion Subsystem (MPS) consists of the descent propulsion section (DPS) (Descent Propulsion Section - Component Location Diagram figure 2. 3-1) and the ascent propulsion section (APS) . Each section is complete and independent of the other and consists of a liquid-propellant rocket engine with its own propellant storage, pressurization, f and feed components . The DPS provides the thrust to control descent to the lunar surface. The APS provides the thrust for ascent from the lunar surface. In case of mission abort, the APS and/ or DPS can place the LM into a rendezvous trajectory with the CSM from any point in the descent trajectory; there is a deadman zone immediately above the lunar surface, where abort cannot be accomplished. The choice of engine to be used depends on the cause for abort, on how long the descent engine has been operating, and on the quantity of propellant remaining in the descent stage.

Descent Propulsion Section - Component Location Diagram

Both propulsion sections use identical hypergolic propellants: a 50-50 mixture, by weight, of hydrazine (N2H4) and unsymmetrical dimethylhydrazine (UDMH) as the fuel; nitrogen tetroxide (N204), as the oxidizer. The injection ratio of oxidizer to fuel is approximately 1.6 to 1, by weight.

Basic operation of the two propulsion sections is similar. In each section, gaseous helium forces the propellants from their tanks , through propellant shutoff valves, to the engine injectors. The DPS uses supercritical helium for propellant pressurization; the APS uses ambient gaseous helium. The primary reason for using supercritical helium is the weight saving. Both the descent and ascent engine assemblies consist of a combustion chamber, where the propellants are mixed and burned; an injector that determines the spray pattern of the propellants injected into the combustion chamber; and propellant control valves and orifices that meter, start , and stop propellant flow to the engine upon command. The descent engine , which is larger and produces more thrust than the ascent engine , is throttleable for thrust control and is gimbaled for thrust vector control. The ascent engine is neither throttleable nor gimbaled. Redundancy of vital components in both propulsion sections provides a high reliability factor.

Before starting the descent or ascent engine, proper propellant settling must be established. This is accomplished by moving the LM in the +X-direction to cause the propellants to settle at the bottom of the tanks . As the propellants are consumed, tank ullage increases and more propellant settling time is required for each subsequent engine start. The +X-translation is accomplished by operating the Reaction Control Subsystem (RCS) downward-firing thrust chamber assemblies (TCA's). Two or all four downward-firing TCA's can be selected, depending upon whether RCS propellant conservation (two TCA's) or a shorter RCS firing time (four TCA's) is the major consideration.

2.3.2 DESCENT PROPULSION SECTION INTERFACES. (See Descent Propulsion Section - Interface Diagram figure 2. 3-2 . )

The DPS receives 28-volt d-c and 115-volt a-c primary power through the Commander's and LM Pilot's buses of the Electrical Power Subsystem (EPS). The outputs of the DPS pressure and temperature transducers and liquid-level sensors are processed in the Instrumentation Subsystem (IS) and are transmitted via the Communications Subsystem (CS) to MSFN: The IS also processes the DPS caution and warning and display signals. The Explosive Devices Subsystem (EDS) opens explosive valves in the DPS to enable propellant tank pressurization and venting.

Descent Propulsion Section - Interface Diagram

The Guidance , Navigation and Control Subsystem (GN&CS) issues automatic on and off commands, gimbal drive actuator commands, and thrust level commands to the descent engine. The automatic on and off commands and thrust level commands can be overriden manually. Descent engine arming and ignition are controlled by automatic guidance equipment, or by the astronauts through the stabilization and control (S&C) control assembly and the descent engine control assembly (DECA). A descent engine arm signal is sent to the S&C control assembly when an astronaut sets the ENG ARM switch (panel 1) to DES or when he presses the ABORT pushbutton (panel 1) preparatory to starting a mission abort program, using the descent engine. Engine-on signals from the LM guidance computer (LGC) or abort guidance section (AGS) are sent to the DECA through the S&C control assembly. The DECA turns the descent engine on upon receiving the arm and the engine-on signals. If DECA power fails, the DES ENG CMD OVRD switch (panel 3) , in the ON position, will supply an alternate voltage source to keep the engine firing. The engine remains on until the engine-off command is received from the automatic guidance equipment. The astronauts can also generate engine on and off commands manually; these commands override the automatic commands . A manual start is accomplished (after propellant tank pressurization with ambient helium) by arming the descent engine and pressing the START pushbutton (panels 5). Either astronaut can shut off the descent engine by pressing his STOP pushbutton (panels 5 and 6) or by pressing the ABORT STAGE pushbutton (panel 1) . An abort-stage command results in immediate descent engine shutdown, automatically followed by ascent propellant tank pressurization, and enabling of circuitry for stage separation and ascent engine firing. Stage separation and ascent engine firing occurs when the ascent engine-on command is issued.

Descent engine throttling is controlled by the LGC or the astronauts . The throttling-range limitations are from minimum thrust (approximately 10% of 10,500 pounds) to approximately 65% and full throttle (approximately 92.5%). The range between 65% and 92.5% is a transient region that cannot be used for extended periods because excessive engine erosion occurs in this zone . Under normal conditions, the engine cannot be operated in the transient region because automatic throttle commands above 65% automatically produce a full throttle output. Only in case of malfunction can inadvertent throttling occur in the transient region, in which case manual correction must be made. Automatic throttle increase and decrease signals from the LGC are sent to an integrating counter in the DECA. The analog output of the DECA controls descent engine thrust. In the automatic mode, the thrust/translation controller assemblies (TTCA's) can be used by the astronauts to increase descent engine throttle (overriding the automatic throttle command); the TTCA' s cannot be used, however, to decrease the throttle command. (Refer to paragraph 2.3.3.3.1) In the manual throttle mode, the astronauts have complete control over descent engine thrust.

The primary guidance and navigation section (PGNS) of the GN & CS, or the AGS, automatically controls descent engine gimbal trim, to compensate for center-of-gravity offsets during descent engine firing. In PGNS operation, the LGC sends trim on and off signals in two directions, for each gimbal axis, to the DECA. These signals operate power control circuitry, which drives the two gimbal drive actuators (GDA's). In AGS operation, Y- and Z-axis error signals from the attitude and translation control assembly (ATCA) are sent to the DECA to drive the GDA"s. The GDA's tilt the descent engine along the Y-axis and Z -axis a maximum of 6° from the center position. The ENG GMBL switch (panel 3) permits removing GDA power to interrupt the tilt capability if the ENG GMBL caution light (panel 2) goes on, indicating a malfunction.

The DPS modes of operation are discussed in detail in paragraph 2.1.3.5. The control circuitry is shown in figure 2.1.18.

2.3.3 DESCENT PROPULSION SECTION FUNCTIONAL DESCRIPTION

The DPS consists of an ambient helium bottle and a cryogenic helium storage vessel with associated helium pressurization components; two fuel and two oxidizer tanks with associated feed components; and a pressure-fed, ablative, throttleable rocket engine. The engine can be shut down and restarted, within operational limitations and restrictions (paragraph 2.3.6), as required by the mission. At the fixed full-throttle position, the engine develops a nominal thrust of 9,870 pounds; it can also be operated within a nominal range of 1,050 to 6,800 pounds of thrust.

The engine is mounted in the center compartment of the descent stage cruciform; it is suspended, at the throat of the combustion chamber, on a gimbal ring that is part of the engine assembly. The gimbal ring is pivoted in the descent stage structure, along an axis normal to that of the engine pivots . The engine can be tilted up to +6° or -6°, by means of the GDA's, to ensure that the thrust vector passes through the LM center of gravity.

Functionally, the DPS can be subdivided into a pressurization section, a propellant feed section, and an engine assembly.

2.3.3.1 Pressurization Section. (See Descent Propulsion Section - Simplified Functional Flow Diagram figure 2. 3-3. )

Before earth launch, the propellant tanks are only partly pressurized, so that the tanks will be maintained within a safe pressure level under the temperature changes that can occur between the time the tanks are loaded and launch. Before initial engine start, the ullage space in each propellant tank requires additional pressurization. This initial pressurization (prepressurization) is accomplished with ambient helium. (Supercritical helium cannot be used because the helium circulating through the fuel/helium heat exchanger may freeze the fuel before fuel flow is established.) A pressure transducer at the outlet port of the ambient helium bottle supplies a signal through the HELIUM MON selector switch (panel 1), when set to AMB PRESS, to the HELIUM indicator (panel 1) to enable the astronauts to check the status of the bottle before initial engine start. The propellant tanks are prepressurized by opening explosive valves in the ambient helium line and in the lines leading to the fuel and oxidizer tanks . The valve in the ambient helium line prevents helium flow from the storage bottle before prepressurization. The compatibility valves in the lines leading to the fuel and oxidizer tanks prevent propellant vapors from degrading the upstream components due to prolonged exposure before pressurization.

After setting the MASTER ARM switch (panel 8) to ON, the DES PRPLNT ISOL VLV switch (panel 8) is set to FIRE to open the fuel and oxidizer compatibility valves . The DES START He PRESS switch (panel 8) is then set to FIRE, opening the ambient helium isolation explosive valve. Ambient helium flows from the storage bottle through the open explosive valve and through a filter, where debris from the explosive valve is trapped. The ambient helium enters the main pressurization line downstream of the normally closed helium isolation solenoid valve and flows through the secondary pressure regulator where pressure is reduced to approximately 245 psi. The regulated ambient helium then enters the propellant tanks to provide the normal ullage pressure.

Descent Propulsion Section - Simplified Functional Flow Diagram

After pressurization with ambient helium, the supercritical helium from the cryogenic storage vessel is used to maintain ullage pressure. Supercritical helium is stored at a density approximately eight times that of ambient helium. Because heat transfer from the outside to the inside of the cryogenic storage vessel causes a gradual increase in pressure (approximately 6 to 10 psi per hour , depending on ambient conditions). Only a component malfunction, a significant time slippage , or a change in the predetermined burn profile can cause the supercritical helium to approach, or exceed, the limits of the pressure/time envelope. An out-of-limit 1 condition may result in rupture of the dual burst disks which may cause the descent engine to operate in a blowdown mode. (The period of engine operation in the blowdown mode depends upon the amount of ullage volume present.)

The cryogenic storage vessel is isolated by an explosive valve , which is fired automatically after an engine-on command has been given. With the ENG ARM switch (panel 1) set to DES, the engineon command can be given manually by pressing the START pushbutton (panel 5}, or it can be given automatically by the LGC or AGS. In either case, the MASTER ARM switch must be in the ON position to fire the explosive valve . A delay circuit causes a 1.3-second delay between opening of the propellant shutoff valves and firing of the supercritical helium isolation explosive valve. This time delay prevents the supercritical helium from entering the fuel/helium heat exchanger until fuel flow is established so that freezing of the fuel in the heat exchanger cannot occur .

The supercritical helium initially passes through the first loop of the two-pass fuel/helium heat exchanger. Here it absorbs heat from the fuel that is routed from the fuel tanks through the heat exchanger, before ultimate delivery to the engine. The helium is warmed to approximately -200° F and routed back through the helium/helium heat exchanger inside the cryogenic helium storage vessel. The -200° F helium transfers heat to the remaining supercritical helium in the vessel, causing an increase in pressure in the vessel that ensures continuous expulsion of helium throughout the entire period of operation . After passing through the helium/helium heat exchanger, where it is cooled' to approximately -300° F, the helium is routed back through the second loop of the fuel/helium heat exchanger and heated to approximately +35° F before delivery to the pressure regulators.

With the supercritical helium pressurization system operating, the pressure in the cryogenic helium storage vessel varies between 400 and 1,750 psia. The pressure is monitored on the HELIUM indicator when the HELIUM MON selector switch is set to SUPCRIT PRESS. The cryogenic storage vessel is protected against overpressurization by a dual burst disk assembly . If an excessive heat transfer through the vessel wall increases the internal pressure above approximately 1,900 psid, the burst disks rupture and the entire helium supply is lost. A normally open vent relief valve between the two burst disks protects against back pressurization of the upstream burst disk if it develops a small leak. If a large leak develops, the vent relief valve closes and the downstream burst disk protects the storage vessel. A thrust neutralizer at the outlet of the downstream burst disk prevents generation of unidirectional thrust if the burst disks rupture.

Downstream of the fuel/helium heat exchanger , the helium flow continues through a filter that traps debris from the explosive valve, then the pressurization l ine divides into two parallel legs. A normally open, latching solenoid valve and a pressure regulator are in series in the primary leg; a • normally closed, latching solenoid valve is in series with a pressure regulator in the secondary leg. The pressure of the helium flowing through the primary leg is reduced by the pressure regulator to the nominal pressure (245 psia) required to pressurize the propellant tanks . If this regulator fails open or closed, pressure at the helium manifold increases or decreases accordingly beyond acceptable limits (rises above 260 psia or drops below 220 psia) and the DES REG warning light (panel 1) goes on. When a caution or warning light goes on, a signal is routed from the caution and warning electronics assembly. (CWEA) in the IS to light the MASTER ALARM pushbutton/lights {panels 1 and 2} and to provide a 3-kc tone in the astronaut headsets . Pressing either MASTER ALARM pushbutton turns off both lights and terminates the tone, but has no effect on the caution or warning light. (The DES REG warning light is inhibited before initial descent engine arming. It will go off when normal pressure is restored when the CWEA circuit breaker is cycled, or when the ascent and descent stages separate.) Under regulator failure conditions , the astronauts must close the solenoid valve in the malfunctioning leg and open the solenoid valve in the redundant leg , to restore normal propellant tank pressurization. The normally open (primary) solenoid valve is closed by momentarily setting the DESCENT He REG 1 switch (panel 1) to CLOSE; the DESCENT He REG 1 talkback above the switch then provides a barber-pole display. The normally closed (secondary) solenoid valve is opened by momentarily setting the DESCENT He REG 2 switch to OPEN; the DESCENT He REG 2 talkback above the switch then provides a gray display. (Both solenoid valves may be closed during the coast periods of descent, to prevent inadvertent tank overpressurization due to possible helium leakage through the pressure regulators and to inhibit leaks downstream of the latching valves.)

The primary and secondary helium flow paths merge downstream of the regulators to form a common helium pressurization manifold. Transducers monitor the manifold pressure; they provide continuous telemetry signals to MSFN, and signals that cause the DES REG warning light to go on when the sensed pressure exceeds 260 psia or drops below 220 psia. The manifold routes the helium into two flow paths: one path leads to the oxidizer tanks; the other, to the fuel tanks. Each path has a quadruple check valve assembly in a ser1 s -parallel arrangement. The quadruple check valves isolate the upstream components from the corrosive propellant vapors and prevent hypergolic action, as a result of backflow from the propellant tanks, in the helium pressurization manifold. After passing through the compatibility explosive valves, the helium flows into the top of the fuel and oxidizer tanks. Diffusers at the top of the tanks uniformly distribute the helium throughout the ullage space. Helium crossover lines maintain a balanced ullage pressure in the tanks containing the same propellants.

Immediately upstream and in parallel with the propellant tanks, each helium flow path contains a relief valve assembly to protect the propellant tanks against overpressurization. The assemblies (a burst disk in series with a relief valve) vent pressure in excess of approximately 275 psia and reseal the flow paths after overpressurization is relieved (254 psia). Thrust neutralizers eliminate unidirectional thrust generated by the escaping gas. To prevent leakage through single point relief valves during normal operation, the burst disks are located upstream of the relief valves. The burst disks rupture at a pressure between 260 and 275 psi; the relief valves open fully at 275 psi to pass the entire helium flow from a failed-open regulator preventing damage to the propellant tanks.

Two vent lines, in parallel with the relief valve assemblies, include an explosive valve in series with a normally open solenoid valve for each propellant tank. The vent lines are intended for planned depressurization of the tanks after lunar landing, when temperature rise of the supercritical helium and heat soak-back from the engine (after shutoff) causes pressure buildup in the tanks. The planned venting arrangement protects the astronauts against untimely venting of the tanks through the relief valve assemblies. The fumes are vented overboard, through the relief valve thrust neutralizers at the fuel and oxidizer pressurization line vents. (See Descent Propulsion Section - Component Location Diagram figure 2. 3 -1.) If the helium pressurization line is open, the supercritical helium in the cryogenic storage vessel will be vented together with the propellant tanks. The supercritical helium will vent rapidly until pressure drops to approximately 350 psia, then the pressure remaining in the cryogenic storage vessel will decrease with the decreasing propellant tank pressures. To open the vent lines, the MASTER ARM switch is set to ON and the DES VENT switch (panel 8) is set momentarily to FIRE , opening both explosive vent valves simultaneously. The MASTER ARM switch is then set to OFF. Venting of the lines is monitored by setting the PRPLNT TEMP/PRESS MON switch (panel 1) to DES 1 and the HELIUM MON selector switch (panel 1) to SUPCRIT PRESS. When the OXID PRESS indicator indicates less than 20 psia , the HELIUM MON selector switch is set to OFF, the OXID VENT switch is set to CLOSE, and the OXID VENT talkback will change to a barber-pole display. When the FUEL PRESS indicator indicates less than 8 psia , the FUEL VENT switch is set to CLOSE , causing the FUEL VENT talkback to provide a barber-pole display.

2.3.3.2 Propellant Feed Section. (See Descent Propulsion Section - Simplified Functional Flow Diagram figure 2. 3 -3. )

Each pair of propellant tanks (containing like propellants) is manifolded into a common delivery line. Balanced propellant flow is maintained by trim orifices in all propellant lines downstream of the tanks.

Helium pressure in the propellant tanks is monitored on the FUEL and OXID PRESS indicators (panel 1), propellant temperature in the tanks is monitored on the FUE L and OXID TEMP indicators. The PRPLNT TEMP/PRESS MON switch selects the set of fuel and oxidizer tanks (No. 1 or No. 2) for monitoring. Each propellant tank has its own temperature transducer to supply temperature signals to the indicator. One pressure transducer in the fuel pressurization line and one in the oxidizer pressurization line supply pressure signals to the indicators. Therefore, the pressure reading remains constant regardless of whether tank No. 1 or 2 monitored. Propellant quantity remaining in the tanks is monitored on the OXIDIZER and FUEL QUANTITY indicators (panel 1). The PRPLNT QTY MON switch selects the set of fuel and oxidizer tanks (No. 1 or 2) for monitoring.

Pressurized helium, acting on the surface of the propellant, forces the fuel and oxidizer into the delivery lines through a propellant retention device that maintains the propellant in the delivery lines during negative-g acceleration (up to acceleration in excess of -2g). The oxidizer is piped directly to the engine assembly; the fuel circulates through the fuel/helium heat exchanger before it is routed to the engine assembly. A small bypass permits some fuel to reach the engine without flowing through the heat exchanger. This protects against a pressure buildup should the fuel in the heat exchanger be frozen. Each delivery line contains a trim orifice and a filter. The trim orifices provide engine interface pressure of approximately 222 psia at full throttle position for proper propellant use. The filters prevent debris, originating at the explosive valves or in the propellant tanks, from contaminating downstream components.

2.3.3.2.1 Propellant Quantity Gaging System. (See Propellant Quantity Gaging System - Simplified Functional Block Diagram figure 2. 3-4. )

The propellant quantity gaging system (PQGS) enables the astronauts to continuously monitor I the quantity of, propellants remaining in the four tanks. The PQGS is of the capacitance type. It consists of four quantity-sensing probes with low-level sensors, a control unit, two QUANTITY indicators, the PRPLNT QTY MON switch , and the DES QTY warning light. During a lunar-landing mission, the PQGS will be turned on approximately 10 seconds before engine ignition and shut off approximately 10 seconds after engine shutdown. The continuous PQGS power-on time is limited to 45 minutes. This limitation safeguards the thermal capability of the electronic components which, if exceeded, could result in erroneous indications. The PROPUL: PQGS circuit breaker (panel 16) is used to apply or remove PQGS power. The PRPLNT QTY MON switch selects a set of propellant tanks (fuel and oxidizer tanks No. 1 or 2) to be monitored on the FUEL and OXIDIZER QUANTITY indicators. With the PRPLNT QTY MON switch set to OFF, the QUANTITY indicators remain lit; however, the digital readouts on the indicators blank out. With the PRPLNT QTY MON switch set to DES 1 or DES 2 and ttie descent engine shut off, the QUANTITY indicator readings remain stable until a zero-g condition develops, at which time the readings drift and become indeterminate.

Propellant Quantity Gaging System - Simplified Functional Block Diagram

The low- level sensors provide a discrete s ignal to cause the DES QTY warning light to go on when the propellant level in any tank is down to 9.4 inches (equivalent to 5. 6% propellant remaining, or sufficient for 116 seconds of engine burn at hover thrust (approximately 25%)). The MASTER ALARM pushbutton/light and the 3-kc tone are not activated when the DES QTY warning light is energized to prevent distraction of the astronauts during the most critical phase of the lunar landing mission. The PQGS has an estimated uncertainty tolerance of 1.3% of full tank capacity for cabin display and telemetry transmission. This tolerance is reduced to 1% in the 8% to 25% propellant quantity range where the PQGS performs at a higher accuracy.

The quantity-sensing probes are double-walled; one probe is installed to run vertically through the center of each propellant tank. Each probe has a 20-volt d-c input. Varying resistance, caused by propellant consumption, causes the output signal to vary from 5 volts de to zero, in direct proportion to propellant quantity. The output signals, processed through the probe circuitry, are sent directly to the control unit which converts them to provide the following:

Indications of the quantity of propellant remaining in the tanks, on the QUANTITY indicators
Conditioned signals, corresponding to propellant quantities in each of the four tanks, to the telemetry registers
A warning signal when the propellant in any of the four tanks is depleted to a predetermined level.

2.3.3.3 Engine Assembly. (See Descent Engine Assembly - Flow Diagram figures 2. 3-5 and Descent Engine and Head End Assembly 2.3-6. )

Fuel and oxidizer enter the engine assembly through interface flanges on opposite sides of the engine. The fuel line has a tap-off branch (pilot valves actuation line) that leads through two actuator isolation valves (arranged in parallel for redundancy) to the four solenoid-operated pilot valves. The fuel in this line is routed, through the pilot valves, to the actuators, where it is used as actuation fluid to open the propellant shutoff valves. The main fuel and oxidizer flow is routed through respective flow control valves, then each flow path splits into two parallel paths that route the propellants through the redundant propellant shutoff valves. The propellant shutoff valve assemblies are in a series-parallel arrangement. The series redundancy prevents open failure; the parallel redundancy prevents closed failure. The valves open simultaneously to permit propellant flow to the engine while it is operating; they close simultaneously to terminate propellant flow at engine shutdown. At the two upstream fuel shutoff valves, venturis restrict the fuel flow so that the oxidizer reaches the injector between 40 and 50 milliseconds before the fuel. This precludes the possibility of a fuel lead, which would result in rough engine starts. Downstream of the propellant shutoff valves, the parallel paths merge to form one fuel and one oxidizer path. The fuel passes through a final trim orifice and enters the variable-area injector manifold, where a concentric annulus of fuel flow is formed. The oxidizer is routed, through the center element of the injector, to the combustion chamber, where it mixes with the fuel for combustion.

Descent Engine Assembly - Flow Diagram

Descent Engine and Head End Assembly

Before initial engine operation and during engine shutdown, the solenoid-operated actuator isolation valves (pre-valves) are closed to prevent possible fuel loss in the pilot valve actuation line due to leakage at the pilot valves. The actuator isolation valves are opened by setting the ENG ARM switch to DES. This enables the actuation fuel to flow to the pilot valves just before the pilot valves are opened. When the START pushbutton is pressed (or an engine-on command initiated), the four solenoid-operated pilot valves open simultaneously, permitting the actuation fuel to open the propellant shutoff valves, thus routing fuel and oxidizer to the combustion chamber. During the start, the solenoids in the pilot valves unseat the caged balls from the inlet ports and seat them against the overboard vent ports; fuel enters the actuator cavities. The actuator pistons are connected to rack-and-pinion linkages that rotate the balls of the shutoff valves 90° to the open position to permit propellant flow to the injector. The series- ( parallel redundancy in the valve arrangement provides for positive start and shutdown. During shutdown, the solenoids in the pilot valves are deenergized and the vent ports are open. The spring-loaded actuators close the shutoff valves. Residual actuation fuel is vented overboard through four separate lines that lead to vent ports at the bottom of the descent stage. (See Descent Propulsion Section - Component Location Diagram figure 2. 3-1. )

The propellant in the main fuel and oxidizer lines flows through cavitating-venturi-type flow control valves that control propellant flow to the engine below the 65% throttle setting. Transition from cavitation to noncavitation occurs between 70% and 80%. At full throttle, and during momentary I transition through the full throttle to 65% range, engine throttling takes place primarily in the pintle assembly of the injector and in the flow control valves. At approximately 70% of maximum thrust, cavitation commences in the throats of the flow control valves, causing the valves to function as cavitating venturis down to minimum thrust. Once cavitation begins, the propellant-metering function is entirely removed from the injector; flow is controlled entirely by the flow control valves.

The throat area of the flow control valves is regulated by a close-tolerance, contoured, metering pintle that is linked directly to the injector sleeve. The linkage is operated by a single actuator so that movement of the actuator simultaneously adjusts the pintles in the flow control valves and the movable sleeve in the injector. The fuel and oxidizer are injected at velocities and angles compatible with variations in weight flow. At full throttle, engine operation is conventional. As the engine is throttled down, the pintles in the flow control valves are moved to decrease the flow control area in the venturis so that the pressure drop across the valves balances out the differential between engine and injector inlet pressures. At the same time, the injector orifice areas are decreased so that the injection velocities and impingement angles of fuel and oxidizer are maintained at near-optimum condition for combustion effieiency.

The injector consists of a faceplate and fuel manifold assembly with a coaxial oxidizer feed tube and an adjustable (metering) orifice sleeve . Oxidizer flows through the center tube and out between a fixed pintle and the bottom of the sleeve; the fuel orifice is an annular opening between the sleeve contour and the injector faceplate. The fuel flows behind the face of the injector to cool the faceplate . Some fuel, tapped off the fuel manifold is used for barrier cooling of the wall, The fuel is emitted in the form of a thin cyclindrical sheet; the oxidizer, in a series of individual sprays. The oxidizer sprays break up the fuel stream and establish the injector pattern at all thrust settings.

The mechanical linkage that connects the pintle of the flow control valves and the injector sleeve is pivoted about a fulcrum attached to the injector body. Throttling is controlled by the throttle valve actuator, which positions the linkage in response to electrical input signals. At maximum thrust, the actuator positions the linkage to fully open the flow control valves and injector apertures . As commanded thrust is reduced, the actuator reduces the flow at the flow control valves and moves the injector orifice sleeve to reduce the apertures . As the adjustable orifice sleeve moves upward, the area of the propellant orifices increases.

2.3.3.3.1 Descent Engine Control

After the engine is manually armed by setting the ENG ARM switch to DES, it can be fired automatically or manually. Under manual control, the engine can be started and stopped by the Commander by pressing the START and STOP pushbuttons on panel 5; it can be stopped by the LM Pilot by pressing the STOP pushbutton on panel 6. The mode of thrust control is determined by the THR CONT switch (panel 1). When this switch is set to AUTO, engine thrust is controlled by the LGC. When the switch is set to MAN, the Commander's or LM Pilot's TTCA (depending on the setting of the MAN THROT switch, panel 1) controls the engine thrust. In the automatic mode , the TTCA is still operational. It is normally set against its hard (low) stop, where it supplies a 10% thrust command that is summed with the LGC command. resulting in a combined thrust command to the descent engine. For example , if the required thrust is 50%, the LGC commands 40%, which is augmented by the 10% obtained from the TTCA. If the TTCA is moved from the hard stop, it supplies a greater portion of the combined command and the LGC command decreases accordingly. Thus, for the required 50% thrust, the TTCA may now command 20%; the LGC, 30%. If the TTCA is moved to a setting such that it commands more than the required thrust, it overrides the automatic command (the LGC portion becomes zero) and descent engine thrust is determined entirely by the TTCA setting.

The dual-scale CMD THRUS'I and ENG THRUST indicator (panel 1) displays commanded manual or automatic thrust on the CMD scale and actual engine thrust on the ENG scale. The ENG scale: input is derived from a pressure transducer in the combustion chamber, because thrust is proportional to chamber pressure. At full throttle position the ENG scale reads 92 . 5% (actualfull-throttle-position thrust) •I while the CMD scale reads between 92.5% and 100%. At all other throttle settings (10% to 65% throttling range) the ENG and CMD scales normally display identical readings of the actual engine thrust. Display of dissimilar readings indicates that the engine is not following the thrust commands or that transfer from automatic to manual throttle control is in process. As shown in the example given previously, if the TTCA is displaced from the hard stop in the automatic mode, for 50% required thrust, the TTCA may command 20% while the LGC contributes 30%. The ENG scale of the THRUST indicator will read 50%; however, the CMD scale (where LGC command is summed with a 10% bias) will read 40%. As the TTCA is moved to increasing throttle settings, the CMD scale readings decrease. When the CMD scale reading has dropped to 10%, the LGC no longer supplies any portion of the thrust command and the TTCA is in contrCJl. At this point, a smooth transfer from the automatic mode to the manual mode is accomplished by setting the THR CONT switch to MAN. The CMD and ENG scales will now indicate identical readings . (For the preceding example, both pointers will align at 50%. ) Very slight deviations between CMD and ENG scale readings may occur as engine operating time increases. The deviations are due to combustion chamber erosion, which causes chamber pressure to decay slightly.

The T/W (thrust/weight) indicator (panel 1) is used primarily to monitor X-axis acceleration during lunar landing and lift-off. The T/W indicator is a self-contained accelerometer that displays instantaneous X-axis acceleration in lunar-g units (1 lunar g = 5. 23 ft/sec2). Inasmuch as a given throttle setting provides a specific acceleration when the vehicle has a given mass, the T/W indicator can be used a.s backup for the THRUST indicators to monitor engine performance.

2.3.4 DESCENT PROPULSION SECTION MAJOR COMPONENT/FUNCTIONAL DESCRIPTION

2.3.4.1 Explosive Valves. (See Descent Propulsion Section - Explosive Valves, Simplified Functional Diagram figure 2.3-7.)

An ambient helium isolation valve , a supercritical helium isolation valve , a fuel compatibility valve, an oxidizer compatibility valve, a fuel vent valve, and an oxidizer vent valve are the explosive valves used in the DPS. These valves normally are closed; they are controlled by the EDS control and fire circuits and, when fired. fully open and remain open. To prevent valve failure in the closed position, each explosive valve has two cartridges that are fired from redundant systems in the EDS. A cartridge is fired by applying power to the initiator bridgewire for a few milliseconds. The resultant heat fires the initiator, generating gases in the valve explosion chamber at an extremely high rate. The gases drive the valve piston into the valve housing to open the valve by shearing a closure disk and aligning the piston port permanently with the pressure line plumbing.

Descent Propulsion Section - Explosive Valves, Simplified Functional Diagram

2.3.4.2 Cryogenic Helium Storage Vessel

The cryogenic helium storage vessel is double walled; it consists of an inner spherical tank and an outer jacket. The void between the tank and the jacket is filled with aluminized mylar insulation and evacuated to minimize ambient heat transfer into the tank. The vessel has fill and vent ports, a burst disk assembly, and an internal helium/helium heat exchanger. The inner tank is initially vented and loaded with liquid helium at approximately 8° R; the fill sequence is completed by closing the vent and introducing a high-pressure head of gaseous helium. As the high-pressure, low-temperature gas (at approximately 14° R) is introduced, the density and pressure of the stored helium are increased. At the end of pressurization, the density of the stored supercritical helium is approximately 8.7 pounds per I cubic foot, and the final pressure is approximately 80 psi.

The burst disk assembly (Cryogenic Helium Storage Vessel - Burst Disk Assembly Diagram figure 2. 3-7A) prevents hazardous overpressurization within the vessel. It consists of two burst disks in series, with a normally open, low-pressure vent valve between the disks. The burst disks are identical; they burst at a pressure between 1,881 and 1,967 psid to vent the entire supercritical helium supply overboard. A thrust neutralizer at the outlet of the downstream burst disk diverts the escaping gas into opposite directions to prevent unidirectional thrust generation. The vent valve prevents low-pressure buildup between the burst disks if the upstream burst disk leaks slightly. The valve is open at pressures below 150 psia. It closes when the pressure exceeds 150 psia, so that, for faster leaks , the pressure buildup will eventually rupture the downstream burst disk and all super critical helium will be vented.

Cryogenic Helium Storage Vessel - Burst Disk Assembly Diagram

2.3.4.3 Fuel/Helium Heat Exchanger

Fuel is routed directly from the fuel tanks to the two-pass fuel/helium heat exchanger. The heat exchanger transfers heat from the fuel to the supercritical helium, which is warmed to operating I temperature by flowing through the two separate heat exchanger passes. The fuel/helium heat exchanger is of finned tube construction; the first and second helium passages are in parallel crossflow with respect to the fuel. All parts are stainless steel; each fin is lap brazed to an adjacent fin and to the side panels to increase structural rigidity. Helium flows in the tubes and fuel flows in the outer shell across the bundle of staggered, straight tubes.

2.3.4.4 Helium Isolation Solenoid Valves (See Helium Isolation Solenoid Valve, Latched-Open Position Diagram figure 2. 3-8. )

These helium isolation valves are two-coil , latching, solenoid-operated valves that shut off helium flow through one leg of redundant flow lines if the pressure regulator in the leg fails open. The valves are actuated by the DESCENT He REG 1 and REG 2 switches (panel 1). A position indicator switch in each valve feeds a signal to the DESCENT He REG 1 and REG 2 talkbacks to indicate whether the particular valve is closed (barber-pole display) or open (gray display) .

Helium Isolation Solenoid Valve, Latched-Open Position Diagram

The valve poppet is part of the "open" armature and shaft assembly. When the "latch open" coil is energized, the poppet leaves the valve seat as the "open" armature slides into the latched position, overcoming the force of the closing spring. The latch balls become seated in the "open" armature detents to maintain the valve open after the "latch open" coil is deenergized. To close the valve . the latch release coil is energized, causing the latch armature and pin assembly to slide back, overcoming the force of the latch spring. As the pin moves back, the latch balls drop out of the detents in the "open" armature and the force of the closing spring pushes and maintains the poppet firmly against the valve seat. The normally closed, internal bleed valve opens if seepage past the moving valve seals causes a pressure buildup in the valve housing. An arc suppression circuit, consisting of two zener diodes across each coil, eliminates the induced voltage generated when the coil is deenergized.

2.3.4.5 Helium Pressure Regulators (See Descent Propulsion Section - Helium Pressure Regulators Diagram figure 2. 3-9. )

The primary and secondary helium pressure regulators are two separate units contained in a common housing. Each regulator is supplied from a separate helium pressurization line, but both feed into a s ingle outlet manifold. Each pressure regulator consits of a direct-sensing main stage and a pilot stage. The valve in the main stage is controlled by the valve in the pilot stage, which senses small changes in the regulator outlet pressure (Pr) and converts these changes to proportionally large changes in control pressure (Pc). The main stage valve poppet is positioned for varying flow demands by changes in the control pressure acting on the main stage bellows sensor .

Descent Propulsion Section - Helium Pressure Regulators Diagram

A reduction in flow demand causes a rise in the regulator outlet pressure because flow from the regulator exceeds the new downstream demand. The rise in outlet pressure reduces the pilot valve output, thereby reducing flow into the main stage chamber. Because the pilot stage chamber bleeds directly into the regulator outlet line, reduced flow into the chamber causes a proportional reduction in the control pressure, which, in turn, moves the main stage valve poppet toward the closed position. The resultant reduced flow through the main stage valve matches the downstream demand. An increase in the downstream demand causes a reduction in outlet pressure, which tends to open the pilot valve. The resultant increase in control pressure causes the main stage valve poppet to open more to meet the increased downstream demand.

The flow limiter at each regulator inlet restricts maximum flow through the regulator to 19 pounds of helium per minute so that the propellant tanks are protected if the regulator fails open. The filter downstream of the flow limiter prevents particles, which could cause excessive leakage at lockup or regulation malfunction, from reaching the main stage and the pilot valve seats. The control pressure relief valve prevents control pressure from exceeding the regulator outlet pressure by more than 150 to 200 psi, by venting main stage chamber pressure into the outlet line. This relief valve operates only if the pilot valve fails; it does not interfere with normal pressure regulator operation. The overstroke relief valve at the main stage bellows sensor prevents the regulated outlet pressure from overshooting its limit.

2.3.4.6 Relief Valve Assemblies (See Descent Propulsion Section - Relief Valve Assembly figure 2. 3-10. )

The relief valve assemblies are downstream of the burst disk assemblies in the helium pressurization line. The helium enters the relief valve chamber and acts against the relief valve diaphragm, overcoming the force of the relief valve spring and lifting the poppet off its seat. This opens the I helium pressurization line to vent excess helium overboard. A thrust neutralizer at the outlet port prevents generation of unidirectional thrust. The cylindrical filter in the relief valve chamber prevents particles from lodging in the valve seat. When helium pressure drops· below the reseat pressure , the relief valve closes to prevent further helium loss. Before burst disk rupture , the light vent spring keeps the relief valve poppet slightly off its seat. This prevents pressure buildup in the relief valve chamber in case of a slight burst disk leak. If the pressure in the relief valve chamber exceeds approximately 8 psi, the force of the vent spring is overcome and the relief valve poppet closes. Pressure then continues to build up and work against the relief valve diaphragm. When the pressure reaches approximately 260 psi, the relief valve cracks .

Descent Propulsion Section - Relief Valve Assembly

2.3.4.7 Burst Disk Assemblies (See Descent Propulsion Section - Burst Disk Assembly figure 2. 3-11. )

Each helium pressurization line leading to the propellant tanks has a burst disk assembly upstream of the relief valve assembly. The function of the burst disk assembly is to protect the relief valve from corrosion by the propellant. Between 260 and 275 psia system pressure the burst disk will rupture, permitting the relief valve to function. When the inlet pressure reaches the burst pressure range, the coil-spring-loaded disk assembly and the belleville spring inner support assembly, which are connected by the burst disk, move downstream together . The disk assembly seats first and because the burst disk cannot withstand the force on the inner support assembly, it ruptures. The disk assembly is moved upstream by the coil spring, opening the flow passage.

Descent Propulsion Section - Burst Disk Assembly

2.3.4.8 Propellant Storage Tanks .

The propellant supply is c ontained in four cylindrical, spherical-ended titanium tanks of identical size and construction. Two tanks contain fuel; the other two, oxidizer. Each pair of tanks I containing like propellants is interconnected at the top , and all propellant lines downstream of the tanks contain trim orifices, to ensure balanced propellant flow. A diffuser at the helium inlet port (top) of each tank distributes the pressurizing helium uniformly into the tank. An antivortex device in the form of a series of vanes, at each tank outlet, prevents the propellant from swirling into the outlet port, thus precluding inadvertent helium ingestion into the engine. Each tank outlet also has a propellant retention device (negative -g can) that permits unrestricted propellant flow from the tank under normal pressur ization , but blocks reverse propellant flow (from the outlet : me back into the tank) under zero-g or negative-g conditions. This arrangement ensures that helium does not enter the propellant outlet line as a result of a negative -g or zero-g condition or propellant vortexing; it eliminates the possibility of engine malfunction due to helium ingestion.

2.3.4.9 Throttle Valve Actuator (See Throttle Valve Actuator - Block Diagram figure 2. 3-12. )

The throttle valve actuator is a passively redundant, electromechanical, linear-motion servoactuator which moves the throttle linkage in response to an electrical input command. Moving the throttle linkage simultaneously changes the position of the flow control valve pintles and the injector sleeve . Changing the position of the flow control valve pintles varies the amount of fuel and oxidizer metered into the engine and thus changes the magnitude of engine thrust. The throttle valve actuator is located between the fuel and oxidizer flow control valves; its housing is rigidly attached to the engine head end and its output shaft is attached to the throttle linkage .

Throttle Valve Actuator - Block Diagram

The actuator is controlled by three electronic channels , which power three d-c torque motors on a common shaft. The motor shaft supplies the input to the ball screw , which converts rotary motion to the linear motion of the throttle valve actuator output shaft (total excursion is 0. 754 ± 0. 01 inch) . Nonjamming mechanical s tops prevent overtravel of the output shaft in the retracted and extended positions. All mechanical moving parts of the actuator are w ithin a hermetically sealed portion of the unit (pressurized to 0. 25 psia w ith a 9 to 1 mixture of nitrogen and helium) . A diaphragm- type leak indicator in the cover of the hermetically sealed portion of the actuator provides visual evidence of loss of vacuum within the unit. F ive potentiometers are ganged to the torque motor shaft through a s ingle-stage planeJ2 :·v reduction gear (potentiometer drive gears) . Three of these potentiometers supply position feedback information to the three motor amplifier channels , one to each channel. One of the other two potentiometers provides throttle actuator shaft position data for telemetry to MSFN. the other one is a spare.

The automatic and manual command signal input voltages are summed and fed into the nonlinear shaping network to provide the nonlinear position versus command required to nullify other nonlinearities in the engine and provide an overall signal that represents linear thrust versus command voltage. This signal is summed with the voltage output of the position feedback potentiometer. If the two voltages are of equal magnitude, but positive polarity, the error voltage is zero. If they are not equal, a positive or negative error voltage is generated. (The error voltage is positive if the command signal is greater; negative, if the feedback voltage is greater. ) The error voltage is summed with a preset input from the bias network (used to trim the actuator) and is then conditioned and fed into a preamplifier. The preamplifier output is summed with the preamplifier outputs of the other channels to control the power amplifier that drives the torque motor (clockwise or counterclockwise). As the motor rotates the ball screw, the position feedback potentiometer is driven by the planetary gearing. A current-limiting circuit senses motor current and, through feedback, limits the current to 3 amperes per channel under normal conditions. A current-cutback circuit senses motor voltage and, under conditions indicative of power transistor failure, provides an input to the current- limiting circuit, causing it to limit the current to approximately 0.75 ampere per channel.

The redundancy within the throttle valve actuator ensures tha:t failure of any one electrical component will not cause the actuator to fail. Failure of one preamplifier channel will be nullified by the cross coupling and high-gain feedback and will result in a negligible positioning error. Failure of a power amplifier or torque motor circuit is detected by the current-cutback and/ or current-limiting circuits, and any erroneous output is reduced to a level that can be overcome by the other two channels without detrimental effects.

The throttle valve actuator provides a fail-safe system in the event selective malfunctions external to the throttle valve actuator occur. If either the primary 2 8-volt d-c power, the command and reference voltages, or the +15- or -15 volt reference voltage is lost, the throttle valve actuator causes the descent engine to automatically thrust at full throttle position. If 2 8-volt d-e power is lost while the descent engine is firing, the throttle valve actuator cannot sustain control, it will be driven to the full throttle position by the hydraulic loads imposed by the propellant pressures on the flow control valves and the injector. (If there are no hydraulic loads, the throttle valve actuator moves to the position of balanced mechanical loads.) Loss of the - 15-volt reference voltage causes the potentiometer feedback voltage to drop to zero. As a result, only the positive command voltage is in control, and the throttle valve actuator moves to the full throttle position. If the +15-volt reference voltage fails, the failuredetection circuit that monitors this voltage provides a signal that causes the throttle valve actuator to move to the full throttle position. The power source for the command voltage is the +15 volt power supply that supplies the input to the plus side of the throttle valve actuator reference. Therefore, loss of the command voltage also results in loss of the reference voltage, which causes the throttle valve actuator to move to the full throttle position. (Loss of the command voltage alone would cause the throttle valve actuator to move to the minimum thrust position. )

2.3.4.10 Flow Control Valves. (See Flow Control Valve Diagram figure 2. 3-13 . )

The oxidizer and fuel flow control valves are on the side of the engine , immediately downstream of the propellant inlet lines. They are secured to the throttle valve actuator mounting bracket. The flow control valve pintle assemblies are mechanically linked to the throttle valve actuator by a cross beam.

Flow Control Valve Diagram

The flow control valves are cavitating venturis with movable pintle sleeves . Engine throttling is initiated by an electrical signal to the throttle valve actuator, commanding an increase or decrease in engine thrust. Operation of the throttle valve actuator drives the crossbeam to ( a new position, thus changing the position of the pintles in the flow control valves. This axial movement of the pintles decreases or increases the pintle flow areas to control propellant flow rate and thrust. Below an approximate 70% thrust setting, flow through the valves cavitates, and hydraulically uncouples the propellant transfer system (and thereby, the flow rate) from variations in combustion chamber pressure. The movable pintle sleeve, shaft-driven through the valve body, is sealed by two flexible bellows. Shear pins in the bearing sleeves permit the valve bearings to operate in redundant modes (rolling or sliding) to improve reliability. Normally, the bearings roll; however, under an excessive load, the shear pins give to permit the bearings to slide. Both pintle sleeve shafts are attached to a yoke that is driven by the throttle valve actuator.

2.3.4.11 Propellant Shutoff Valve Assemblies

Each of the four propellant shutoff valve assemblies consists of a fuel shutoff valve, an oxidizer shutoff valve, a pilot valve, and a shutoff valve actuator. The shutoff valve actuator and the fuel shutoff valve are in a common housing. The oxidizer shutoff valve is in a separate housing to isolate fuel from oxidizer. It is actuated by a mechanical linkage driven from the fuel shutoff valve. The fuel and oxidizer shutoff valves , downstream of the flow control valves, are exposed to the full propellant pressure before engine firing. Propellant leakage at the valve seats is vented overboard.

2.3.4.11.1 Pilot Valves and Shutoff Valve Actuators

The four solenoid-operated pilot valves control the actuation fluid (fuel) for the fuel and oxidizer shutoff valves. Normally, the pilot valves are closed, the solenoids are deenergized, and the actuation fluid is shut off by action of the spring-loaded solenoid plungers on the ball poppet seated in the lower pilot valve seats. The back side of the actuator chamber is vented to preclude a fuel buildup caused by leakage of the pilot valves, and overpressurization caused by an excessive fuel temperature increase. When the pilot valves are open, all solenoids are energized, and the actuation fluid (at approximately 110 psia) acts against the spring-loaded actuator plunger to open the oxidizer and fuel shutoff valves. At this point, the pilot valve poppets are forced against the upper valve seats, sealing off the vent port. When the electrical signal is removed, the valve poppets return to their lower seat to seal off the actuation fluid. The propellant shutoff valves are closed by the return action of the actuator piston springs, which expels the fuel entrapped in the cylinders and valve passages through the pilot vent port.

2.3.4.11.2 Fuel and Oxidizer Shutoff Valves

The fuel and oxidizer shutoff valves are mechanically linked, fuel-actuated ball valves that are arranged in a series-parallel configuration. The valve housings are made of aluminum alloy. The ball element operates against a spring-loaded Teflon seat to assure positive sealing when the valve is closed. All ball valves are supported by ball bearings. The individual valves are rotated by a rack-and pinion-gear arrangement, which translates the linear displacement of the pistons in the shutoff valve actuators.

2.3.4.12 Variable-Area Injector . (See Variable-Area Injector - Quarter Section Diagram figure 2 . 3-14 . )

The variable-area injector consists of a pintle assembly, drive assembly, and manifold assembly. The pintle assembly introduces the propellant uniformly into the combustion chamber. The drive assembly has a twofold function: first, it serves as a passage for conducting the oxidizer into the pintle assembly; second, it contains the bearing and sealing components that permit accurate positioning of the injector sleeve . The injector sleeve varies the injection area so that near-optimum injector pressure drops and propellant velocities are maintained at each thrust level. The primary function of the manifold ssembly is to distribute the fuel uniformly around the outer surface of the sleeve. Fuel enters the manifold assembly and is passed through a series of distribution plates near the outer diameter of the assembly.

Variable-Area Injector - Quarter Section Diagram

At the center of the manifold, the fuel passes through a series of streamlined hole rings that suppress fuel flow discontinuities. From this point, the fuel is admitted into a narrow passage formed by two parallel conical surfaces (the manifold body and a faceplate). The passage smoothes out remaining gross fuel discontinuities and assists in cooling the injector face. The fuel then passes onto the outer surface of the sleeve, past a fuel-metering lip. The fuel is injected as a hollow cylinder so that it reaches the impingement zone with a uniform circumferential velocity profile and without atomizing, at all flow rates. The oxidizer is injected through a double-slotted sleeve so that it forms a large number of radial filaments. Each filament partially penetrates the fuel cylinder and is enfolded by fuel in such away that, following fast liquid-phase reactions and gas evolution, little preferential separation of oxidizer and fuel can occur. For given propellant densities, overall mixture ratio, and injector geometry, there is a range of propellant injection velocity ratios that result in maximum mixture ratio uniformity throughout the resultant expanding propellant spray. When they occur, the liquid-phase reactions generate gas and vapor that atomize and distribute the remaining liquid oxidizer and fuel uniformly in all directions, resulting in high combustion efficiency. A separate fuel manifold feeds the barrier cooling orifices to minimize engine erosion.

2.3.4.13 Combustion Chamber and Nozzle Extension. (See Descent Engine Combustion Chamber, Nozzle Extension, and Heat Shield Diagram figure 2 . 3- 15. )

The combustion chamber consists of an ablative- cooled chamber section, nozzle throat, and nozzle divergent section. The ablative sections are enclosed in a continuous titanium shell and jacketed in a thermal blanket composed of aluminized nickel foil and glass wool. A seal prevents leakage between the combustion chamber and nozzle extension.

Descent Engine Combustion Chamber, Nozzle Extension, and Heat Shield Diagram

The nozzle extension is a radiation-cooled, c rushable skirt, made of columbium with an aluminide coating. It is attached to the combustion chamber case at a nozzle area ratio of 16 to 1 and I extends to an exit area ratio of 54. 0 to 1.

2.3.4.14 Gimbal Ring and Gimbal Drive Actuators

The gimbal ring is located at the plane of the combustion chamber throat. It consists of a rectangular beam frame and four trunnion subassemblies . The assembly permits up to +6° or -6° roll and pitch trim control of the engine. A description of the two gimbal drive actuators is given in paragraph 2.2.4.6.

2.3.5 DESCENT PROPULSION SECTION PERFORMANCE AND DESIGN DATA

The performance and design data for the DPS are given in table 2.3-1.

Descent Propulsion Section - Performance and Design Data

2.3.6 DESCENT PROPULSION SECTION OPERATIONAL LIMITATIONS AND RESTRICTIONS

The operational limitations and restrictions for the DPS are as follows:

The landing gear must be deployed before descent engine firing. If not deployed, the landing gear would be in the path of the descent engine plume and would be damaged.
Before descent engine starts, the RCS +X-axis thrusters must be fired to settle the propellants under weightless conditions. Unsettled or insufficiently settled propellants will result in rough or erratic starts that could lead to engine failure.
Propellant bulk temperatures before descent engine start must be between +50° and +90°F. If temperature limits are exceeded, the resultant rough combustion may adversely affect engine-component performance, and ullage pressure limitations may be exceeded.
The descent engine must be not operated for prolonged periods in the throttling range of 65% to 92.5%. In this range, operation of the cavitating venturis of the flow control valves becomes unpredictable and may cause an improper fuel oxidizer mixture ratio, which will result in excessive engine erosion and early combustion chamber burn-through.
The DPS must not remain pressurized longer than 3.5 days before the anticipated termination of use. If this limit is exceeded, the pressurization section components may fail to operate because of the corrosive nature of the propellants.
The descent engine combustion chamber must not be subjected to more than 1,100 seconds of engine operation. Exceeding this limitation will cause the engine to operate with a severely charred combustion chamber, possibly resulting in burn-through.

2.3.7 ASCENT PROPULSION SECTION INTERFACES. (See Ascent Propulsion Section - Interface Diagram figure 2. 3- 16. )

The APS receives 28-volt d-c and 115-volt a-c primary power through the Commander 's and LM Pilot's buses of the EPS. The outputs of the APS pressure and temperature transducers, low level sensors, and valve position indicator switches are processed in the IS and are transmitted via the CS to MSFN. The IS also processes the APS caution and warning signals. The EDS opens explosive valves in the APS to enable propellant tank pressurization. Interconnect plumbing between the APS propellant tanks and the RCS thrust chamber assembly feed lines permits the RCS to use APS propellants during certain mission phases, thereby conserving the RCS propellant supply.

Ascent Propulsion Section - Interface Diagram

The GN&CS issues automatic on and off commands to the ascent engine. Ascent engine arming, ignition, and shutdown can be initiated by automatic guidance equipment (PGNS or AGS) or by the astronauts. The automatic and manual commands are sent to the S&C control assemblies, which provide sequential control of LM staging and engine on and off commands. The ascent engine is armed manually by setting the ENG ARM switch to ASC or pressing the ABORT STAGE pushbutton. (Either action automatically shuts off the descent engine if it is firing.) After setting the ENG ARM switch to ASC (and after initial propellant tank pressurization), the ascent engine can be started manually by pressing the START pushbutton and stopped by pressing either STOP pushbutton. In the event of an abort-stage command while the descent engine is firing, the S&C control assemblies provide a time delay before commanding LM staging and ascent engine firing. This delay ensures that the descent engine has stopped thrusting before staging occurs. To stop the ascent engine after an abort-stage start, the ABORT STAGE pushbutton must be reset (pressed a second time) to release the switch. (This procedure is necessary because the ABORT STAGE pushbutton disables the STOP pushbutton. ) The manual commands override the commands issued by the automatic guidance equipment.

The APS modes of operation are discussed in detail in paragraph 2. 1. 3. 5. The control circuitry is shown in the Ascent Engine Assembly - Flow Diagram figure 2.1.19.

Ascent Engine Assembly - Flow Diagram

2.3.8 ASCENT PROPULSION SECTION FUNCTIONAL DESCRIPTION.

The APS consists of a constant-thrust rocket engine that is not gimbaled, two propellant tanks, two helium tanks , and associated propellant feed and helium pressurization components. (See Ascent Propulsion Section - Component Location Diagram figure 2. 3-17. ) The engine is installed in the midsection of the ascent stage; it is canted so that the center line is tilted 1.5° from the X-axis, in the +Z-direction. The engine develops 3,500 pounds of thrust in a vacuum, sufficient to launch the ascent stage from the lunar surface- and place it into a predetermined lunar orbit. The engine can be shut down and restarted, within operational limitations and restrictions (paragraph 2.3.11), as required by the mission.

Ascent Propulsion Section - Component Location Diagram

2.3.8.1 Pressurization Section. (See Ascent Propulsion Section - Simplified Functional Flow Diagram figure 2. 3-18. )

The propellants are pressurized by gaseous helium, supplied from two identical tanks and routed through redundant flow lines into the propellant tanks. The helium is stored at a nominal pressure of 3,050 psia at a nominal temperature of +70°F. A pressure transducer at each tank outlet port is connected to the H ELIUM indicator via the H ELIUM MON selector switch. B efore staging, the pressure transducers also supply a signal to the IS that will cause the ASC PRESS warning light to go on when the pressure in either helium tank is less than 2,773 psia. This alerts the astronauts to identify and isolate the faulty tank. The HELIUM MON selector switch selects the tank pressure to be displayed. When a caution or warning light goes on, a signal is routed from the CWEA to light the MASTER ALARM pushbutton/lights (panels 1 and 2) and to provide a 3-kc tone in the astronaut headsets. Pressing either MASTER ALARM pushbutton turns off both lights and terminates the tone, but has no effect on the caution or warning light. If the ASC PRESS warning light goes on because pressure in either helium tank is less than 2,773 psia, the light goes off upon separation of the descent and ascent stage.

Ascent Propulsion Section - Simplified Functional Flow Diagram

Before initial ascent engine operation, the helium isolation explosive valves prevent the helium from leaving the tanks. These valves can be opened individually or simultaneously. The propellants are not pressurized until shortly before initial ascent engine start. To accomplish initial pressurization, the helium isolation explosive valves and the fuel and oxidizer compatibility explosive valves (upstream of the propellant tanks) are opened simultaneously Normally, propellant pressurization is initiated by setting the ASCENT He SEL switch (panel 8) to FIRE. If the HELIUM indicator shows a leak (zero or decaying pressure) in one of the tanks, the ASCENT He SEL switch is set so that only the explosive valve leading from the nonleaking tank opens, thus preventing helium loss through the leaking tank via the helium interconnect line (downstream of the explosive valves). The MASTER ARM switch is then set to ON and the ASCENT He PRESS switch is set to FIRE and released, firing the explosive valves in the APS. A filter in each helium flow path traps debris from the explosive valve.

Downstream of the filter , each helium flow path has a normally open latching solenoid valve and two series-connected pressure regulators. The downstream regulator is set to a slightly higher output pressure than the upstream regulator; the regulator pair in the primary flow path produces a slightly higher output than the pair in the secondary (redundant) flow path. This arrangement causes lockup of the regulators in the redundant flow path after the propellant tanks are pressurized, while the upstream regulator in the primary flow path maintains the propellant tanks at their normal pressure of 184 psia. If either regulator in the primary flow path fails closed, the regulators in the redundant flow path sense a demand and open to pressurize the propellant tanks. If an upstream regulator fails open, control is obtained through the downstream regulator in the same flow path. Because the downstream regulator normally does not control the output pressure, an open failure of this regulator has no effect. If both regulators in the same flow path fail open, pressure in the helium manifold increases above the acceptable limit of 220 psia, causing the ASC HI REG caution light (panel 2) to go on. This alerts the astronauts to the fact that the failed-open regulators must be identified and the helium isolation solenoid valve in the malfunctioning flow path must be closed so that normal pressure can be restored. The regulator outlet pressure is sensed by redundant pressure transducers that supply inputs to the IS, for telemetry to MSFN. One of these pressure transducers supplies the input signal to the ASC HI REG caution light. Excessive pressure is vented by the relief valve assemblies in parallel with the propellant tanks . The solenoid valve is closed by setting the ASCENT He REG 1 or REG 2 switch (panels 1) to CLOSE; the talkback above the switch will change to a barber-pole display.

The primary and secondary helium flow paths merge downstream of the regulators to form a common helium manifold. The manifold routes the helium into two flow paths: one path leads to the oxidizer tank; the other, to the fuel tank. A quadruple check valve assembly, a series-parallel arrangement in each path, isolates the upstream components from corrosive propellant vapors. The check valves also safeguard against possible hypergolic action in the common manifold resulting from mixing of propellants or fumes flowing back from the propellant tanks. Two parallel compatibHity explosive valves, downstream of each quadruple check valve assembly, seal off the propellant tank inlets , isolating the fuel and oxidizer (liquid and vapor) before initial ascent engine start. This reduces contamination problems involving helium components and prolongs the life of the pressure regulators. The fuel and oxidizer compatibility explosive valves are opened simultaneously with the helium isolation explosive valves before initial engine start. The four compatibility valves are arranged so that two fuel compatibility valves and two oxidizer compatibility valves, in parallel paths, lead to their propellant tanks . One fuel and one oxidizer compatibility valve has dual cartridges, the other two are fired by single cartridges. (See Ascent Propulsion Section - Explosive Valves, Simplified Functional Diagram figure 2. 3-20. )

Ascent Propulsion Section - Explosive Valves, Simplified Functional Diagram

Immediately upstream of the fuel and oxidizer tanks, each helium path contains a burst disk and relief valve assembly to protect the propellant tanks against overpressurization. This assembly vents pressure in excess of approximately 245 psia (relief valve cracking pressure) and reseals the flow path after overpressurization is relieved. A thrust neutralizer eliminates unidirectional thrust generated by the escaping gas. To prevent leakage through a faulty relief valve during normal operation, the burst disk is located upstream of the relief valve. The burst disk ruptures at a pressure slightly below the relief valve setting. The relief valves can pass the entire helium flow from a failed-open pair of regulators, preventing damage to the propellant tanks.

2.3.8.2 Propellant Feed Section. (See Ascent Propulsion Section - Simplified Functional Flow Diagram figure 2. 3-18. )

The APS has one oxidizer tank and one fuel tank. Each tank has a temperature transducer that supplies propellant temperature signals to the FUEL and OXID TEMP indicators, and a pressure transducer that supplies ullage pressure signals to the FUEL and OXID PRESS indicators . (The same TEMP and PRESS indicators are used for the APS and DPS. ) APS data are displayed when the PRPLNT TEMP/PRESS MON switch is set to ASC. A low-level sensor in each propellant tank causes the ASC QTY caution light (panel 2) to go on when the remaining propellant in either tank is sufficient for only 10 seconds of engine operation. The ASC QTY caution light is inhibited when the ascent engine is not operating.

Helium flows into the top of the fuel and oxidizer tanks . Diffusers at the top of the tanks uniformly distribute the helium throughout the ullage space. The outflow from each propellant tank divides into two paths. In the primary path, each propellant flows through a trim orifice to the propellant filter in the engine assembly, and then to the isolation and bipropellant valve assemblies (propellant shutoff valves). The trim orifice provides an engine interface pressure of 170 psia for proper propellant use. The secondary path connects the ascent propellant supply to the RCS. This interconnection, at the normally closed ascent feed solenoid valves (part of the RCS), permits the RCS to burn ascent propellants , providing the APS is pressurized and the ascent or descent engine is operating when the RCS thrusters are fired. A line branches off the RCS interconnect fuel flow path and leads to two parallel actuator isolation solenoid valves. From there, this line routes fuel to the engine pilot valves that actuate the propellant shutoff valves in the engine assembly. The normally closed actuator isolation valves prevent possible fuel loss through a leaking pilot valve before initial engine operation or during engine shutdown. The actuator isolation valves, and the propellant shutoff pilot valves in the engine assembly, are opened and closed simultaneously by engine-on and engine-off commands.

2.3.8.3 Engine Assembly. (See Ascent Engine Assembly - Flow Diagram figure 2. 3-19. )

The ascent engine is a fixed- injector, restartable, bipropellant rocket engine with an ablative combustion chamber, throat, and nozzle extension. Fuel and oxidizer enter the engine assembly through inlet ports at the interface flanges and are routed through the propellant filters and the propellant shutoff valves (isolation and bipropellant valve assemblies) to the injector; a separate fuel path (actuator pressure line) leads to the pilot valves, where fuel pressure actuates the propellant shutoff valves.

Propellant flow to the engine combustion chamber is controlled by a valve package assembly, trim orifices, and an injector assembly. The valve package assembly consists of eight propellant shutoff ball valves that make up the two fuel-and-oxidizer-coupled isolation valve assemblies and the two fuel-and-oxidizer coupled bipropellant valve assemblies, four actuators, and four solenoid-operated pilot valves. Inside the valve package assembly, the fuel and oxidizer passages divide into dual flow paths, with two series ball valves in each flow path. The paths rejoin at the valve package outlet. The propellant shutoff valves are arranged in fuel-oxidizer pairs; each pair is operated from a single crankshaft assembly by an individual fuel-pressure-operated actuator. Shaft seals and vented cavities prevent fuel and oxidizer from coming into contact with each other due to seepage along the shafts.

After the ascent propellants have been pressurized, the ascent engine can be started manually by setting the ENG ARM switch to ASC and by pressing the START pushbutton. Automatic starts are initiated by LGC or AGS engine-on commands. At engine start, the two actuator isolation valves in the propellant feed section, and the four pilot valves, are opened simultaneously, routing fuel into the actuator feed line and to the four pilot valves. When the solenoids of the pilot valves are energized, the pilot valve spools slide away from the fuel inlet ports and block the overboard vent ports. Fuel enters the actuator chambers and extends the actuator pistons, cranking the propellant shutoff valves 90° to the fully open position. The propellants now pass through the shutoff valves and trim orifices directly to the injector. The orifices determine the thrust level of the engine and the mixture ratio of the propellants by trimming the pressure differentials of the fuel and oxidizer. The physical characteristics of the injector establish an oxidizer lead of between 40 and 50 milliseconds. This precludes the possibility of fuel lead, which would result iu rough engine starts.

At engine cutoff, the pilot valve solenoids are deenergized, opening the actuator ports to the overboard vents so that residual fuel in the actuators is vented into space. (See Ascent Propulsion Section - Component Location Diagram figure 2. 3-17 for vent locations.) With the actuation fuel pressure removed . the actuator pistons are moved back by spring pressure, causing the propellant shutoff valves to turn 90° to the closed position.

The ascent engine assembly has transducers and valve position indicator switches for sensing fuel and oxidizer inlet pressures, thrust chamber pressure, and propellant shutoff valve positions. The transducer outputs are converted to telemetry data in the IS. These data are transmitted to MSFN, where they are used to monitor the performance of the ascent engine.

2.3.9 ASCENT PROPULSION SECTION MAJOR COMPONENT/FUNCTIONAL DESCRIPTION

2.3.9.1 Explosive Valves. (See Ascent Propulsion Section - Explosive Valves, Simplified Functional Diagram figure 2.3 -20. )

A high-pressure helium isolation valve at each helium tank outlet, two low-pressure fuel compatibility valves (in parallel) , and two low-pressure oxidizer compatibility valves (in parallel), are the explosive valves used in the APS. Normally, all helium isolation valves and compatibility valves are opened simultaneously by a command from the EDS control and fire circuits (or from the ABORT STAGE pushbutton on panels 1). The ASC He SEL switch (normally set to BOTH) permits isolating a defective helium tank by setting the switch to TANK 1 or TANK 2, thereby firing either helium isolation valve individually. The ASC He SEL switch determines which helium isolation valves will be opened, regardless of whether manual, automatic, or abort stage fire commands are issued. To prevent valve failure in the closed position, the two high-pressure helium isolation valves have dual cartridges that are fired simultaneously by redundant systems in the EDS. The fuel and oxidizer compatibility valves are in redundant parallel flow paths. To ensure propellant tank pressurization, one fuel valve and one oxidizer valve are fired by dual cartridges, the other two compatibility valves are fired by single cartridges. A cartridge is fired by applying power to the initiator bridgewire for a few milliseconds . The resultant heat fires the initiator, generating gases in the valve explosion chamber at an extremely high rate. The gases drive the valve piston into the valve housing to open the valve by shearing a closure disk and aligning the piston port permanently with the pressure line plumbing.

2.3.9.2 Helium Isolation Solenoid Valves. (See Helium Isolation Solenoid Valve, Latched-Open Position Diagram figure 2.3-8. )

The helium isolation solenoid valves in the APS are identical with those in the DPS except that they are actuated by the ASCENT He REG 1 and REG 2 switches and their position indicator switch signal is fed to the ASCENT He REG 1 and REG 2 talkbacks . (Refer to paragraph 2.3.4.4 for a description of the helium isolation solenoid valves. )

2.3.9.3 Helium Pressure Regulator Assemblies . (See Ascent Propulsion Section - Helium Pressure Regulator Assembly Diagram figure 2. 3-21. )

Each helium pressure regulator assembly consists of two individual pressure regulators connected in series. The downstream regulator functions in the same manner as the upstream regulator; however, it is set to produce a higher outlet pressure so that it becomes a secondary unit that will only be in control if the upstream regulator (primary unit) fails open.

Ascent Propulsion Section - Helium Pressure Regulator Assembly Diagram

Each pressure regulator unit consists of a direct-sensing main stage and a pilot stage. The valve in the main stage is controlled by the valve in the pilot stage which senses small changes in the regulator outlet pressure (Pr) and converts these changes to proportionally large changes in control pressure (Pc). The main stage valve poppet is positioned for varing flow demands by changes in the control pressure acting on the main stage piston. A reduction in flow demand causes a rise in the regulator outlet pressure because flow from the regulator exceeds the new downstream demand. The rise in outlet pressure decreases the pilot valve output, thereby reducing flow into the main stage chamber. Because the main stage chamber bleeds directly into the regulator outlet line (through a fixed orifice), reduced flow into the chamber causes a proportional reduction in the control pressure, which, in turn, moves the main stage valve poppet toward the closed position. The resultant reduced flow through the main stage valve matches the downstream flow demand. An increase in the downstream demand causes a reduction in outlet pressure which tends to open the pilot valve. The resultant increase in control pressure causes the main stage valve poppet to open, thus meeting the increased downstream demand.

The flow limiter at the outlet of the main stage valve of the secondary unit restricts maximum flow through the regulator assembly to 5. 5 pounds of helium per minute, so that the propellant tanks are protected if the regulator fails open. The filter at the inlet of the primary unit prevents particles, which could cause excessive leakage at lockup, from reaching the regulator assembly.

2.3.9.4 Relief Valve Assemblies. (See Ascent Propulsion Section - Relief Valve Assembly Diagram figure 2. 3-22 . )

Each helium pressurization line leading to the propellant tanks has a relief valve assembly that consists of a burst disk in series with a relief valve. The redundancy within each relief valve assembly ensures against helium loss during normal operation and protects the propellant tanks against inadvertent overpressurization. The burst disk is held in place by the automatic initiator assembly. When pressure in the pressurization line exceeds approximately 226 psi, the tensile member breaks, causing the helium pressure and the force of the initiator spring to rupture the burst disk. The helium then enters the relief valve chamber and acts against the relief valve piston, overcoming the main poppet spring, and lifting the poppet off its seat. This opens the helium pressurization line to vent the excess helium overboard. A thrust neutralizer at the oulet port prevents generation of unidirectional thrust. A filter at the inlet to the relief valve chamber prevents particles from lodging in the valve seat. When helium pressure drops below the reseat pressure, the relief valve poppet closes to prevent further helium loss. Before burst disk rupture, the light bleed valve spring keeps the bleed poppet slightly off its seat. This permits a low pressure to be vented and prevents pressure buildup in the relief valve chamber in case of a slight burst disk leak. If the leakage pressure builds up, the force of the bleed valve spring is overcome and the bleed poppet closes. Pressure then continues to build up and work against the relief valve piston. When the pressure reaches approximately 245 psi, the relief valve cracks.

Ascent Propulsion Section - Relief Valve Assembly Diagram

2.3.9.5 Propellant Storage Tanks

The propellant supply is contained in two spherical titanium tanks. The tanks are of identical size and construction. On tank contains fuel; the other, oxidizer. A helium diffuser at the inlet port of each tank distributes the pressurizing helium uniformly into the tank. An antivortex device (a cruciform at each tank outlet) prevents the propellant from swiriling into the outlet port, precluding helium ingestion into the engine. Each tank outlet also has a propellant retention device that permits unrestricted propellant flow from the tank under normal pressurization, but blocks reverse propellant flow (from the outlet line back into the tank) under zero-g or negative-g (better than -2g) conditions. This arrangement ensures that helium does not enter the propellant outlet line while the engine is not firing, thus it eliminates the possibility of engine malfunction due to helium ingestion. A low-level sensor in each tank (approximately 4. 4 inches above the tank bottom) supplies a discrete signal that I causes the ASC QTY caution light to go on when the propellant remaining in either tank is sufficient for approximately 10 seconds of burn time (48 pounds of fuel, 69 pounds of oxidizer minimum).

2.3.9.6 Valve Package Assembly. (See Ascent Engine Assembly Diagram figure 2. 3-23 . )

At the propellant feed section/engine assembly interface, the oxidizer and fuel lines lead into the valve package assembly. The individual valves that make up the valve package assembly are in a series-parallel arrangement to provide redundant propellant flow paths and shutoff capability. The valve package assembly consists of two bipropellant valve assemblies, two isolation valve assemblies, and four solenoid-operated pilot valve and actuator assemblies. Each bipropellant valve assembly and each isolation valve assembly consists of one fuel shutoff valve and one oxidizer shutoff valve . These are ball valves that are operated by a common shaft, which is connected to its respective pilot valve and actuator assembly. Shaft seals and vented cavities prevent the propellants from coming into contact with each other . Separate overboard vent manifold assemblies drain the fuel and oxidizer that leaks past the valve seals , and the actuation fluid (fuel in the actuators upon pilot valve closing), overboard. (See Ascent Propulsion Section - Component Location Diagram figure 2. 3-17. ) The eight shutoff valves open simultaneously to permit propellant flow to the engine while it is operating; they close simultaneously to terminate propellant flow at engine shutdown. The two isolation valve assemblies are exposed to the full pressure of the propellant feed system. The four nonlatching, solenoid- operated pilot valves control the actuation fluid (fuel) for the isolation and bipropellant valve assemblies. Normally, the solenoids are deenergized, and the actuation fuel is shut off by action of the spring- loaded spools blocking the inlet ports. The back sides of the actuator chambers are vented to eliminate fuel buildup due to leakage. When the solenoids are energized, all pilot valves are opened and the actuation fuel enters the actuators where it acts against the spring-loaded actuator piston to open the shutoff (ball) valves. At this point, the pilot valve spools seal off the vent port. When the electrical signal is removed, spring action forces the valve spools to seal off the actuation fuel. The propellant shutoff valves are closed by the return action of the actuator piston springs, and the actuation fuel trapped in the actuator chamber and valve passages is expelled through the pilot valve vent ports.

Ascent Engine Assembly Diagram

2.3.9.7 Injector Assembly

The injector assembly consists of the propellant inlet lines, a fuel manifold , a fuel reservoir chamber, an oxidizer manifold , and an injector orifice plate assembly. It takes longer to fill the fuel manifold and reservoir chamber assembly. Consequently, the oxidizer reaches the combustion chamber approximately 50 milliseconds before the fuel, resulting in smooth engine starts. The injector orifice plate assembly is of the fixed-orifice type, which uses a baffle and a series of perimeter slots (acoustic cavities) for damping induced combustion disturbances. The baffle is Y-shaped, with a 120° angle between each blade. The baffle is cooled by the propellants, which subsequently enter the combustion chamber through orifices on the baffle blades. The injector assembly face is divided into two combustion zones: primary and baffle. The primary zone uses impinging doublets (one fuel and one oxidizer), which are spaced in concentric radial rings on the injector assembly face. The baffle zone (1.75 inches downstream from the injector face) uses impinging doublets placed at an angle to the injector face radius. Film-cooling of the combustion chamber wall is achieved by injecting fuel through orifices spaced around the perimeter of the injector. Parallel orifices inject fuel parallel to the engine centerline; canted orifices inject fuel against the chamber wall to form a film of fuel. The nominal propellant temperature is +70° F as it enters the injector. The temperature range is +50°F to +90°F; the fuel temperature is within 10°F of the oxidizer temperature. The temperature range at engine start may be +40° to +500°F (restart at peak heat soakback).

2.3.9.8 Combustion Chamber Assembly.

The combustion chamber assembly consists of an engine case and mount assembly and a plastic assembly, which includes the nozzle extension. The engine case and mount assembly is bonded and locked to the plastic assembly to form an integral unit. The plastic assembly provides ablative cooling for the combustion chamber; it consists of the chamber ablative material, the chamber insulator, the nozzle extension ablative material, and a structural filament winding. The chamber ablative material extends from the inj ector to an expansion ratio of 4. 6. The chamber insulator, between the ablative material and the case, maintains the chamb er skin temperature within design requirements. The ablative material of the nozzle extension extends from the expansion ratio of 4.6 to 45.6 (exit plane) and provides ablative cooling in this region. The structural filament winding provides the structural support for the plastic assembly and ties the chamber and nozzle extension sections together.

2.3.10 ASCENT PROPULSION SECTION PERFORMANCE AND DESIGN DATA.

The performance and design data for the APS are given in Ascent Propulsion Section - Performance and Design Data table 2. 3-2.

Ascent Propulsion Section - Performance and Design Data

2.3.11 ASCENT PROPULSION SECTION OPERATIONAL LIMITATIONS AND RESTRICTIONS

The operational limitations and restrictions for the APS are as follows:

Propellant tank pressure before pressurization must be between 62 and 205 psia. If these limits are exceeded during dynamic loading, structural failure in the propellant tanks may result, causing loss of the LM.
Propellant bulk temperature before ascent engine start must be between +50° and +90°F. If the temperature limits are exceeded, the engine performance will be degraded.
Before ascent engine starts (except FITH with DPS burn), the RCS +X-axis thrusters must be fired to establish proper propellant tank ullage, to settle the propellants under weightless conditions, and to prevent helium from entering the RCS interconnect lines. Unsettled or insufficiently settled propellants may result in rough or erratic starts that could lead to engine failure.
The APS must not remain pressurized longer than 24 hours before anticipated termination of use. If this limit is exceeded, the pressure regulator assemblies will exceed their qualified propellant exposure time.
In the blow-down mode, it is not desirable that the ascent engine be operated when chamber pressure has decayed to less than 112 psia. At a pressure of less than 114 psia, firing should be terminated unless engine operation is critical to mission success.