LM (DESCENT) SYSTEMSThe whole LM concept was divided into 12 (sub)systems which were
|LM Control and hardware was divided into 12 subsystems. This shows their interaction.|
- GN&CS - Gudance, Navigation and Control
- RS - Radar
- MPS - Main Propulsion
- RCS - Reaction Contro
- EPS - Electrical Power
- ECS - Environmental Control
- CS - Communications
- EDS - Explosive Devices
- IS - Instrumentation
- LS - Lighting
- CPE - Crew Personal Equipment
- CDS - Controls and Displays
|The first Moon landing hit the headlines summer 1969|
The success of the LM powered descent depended on the smooth interaction of key systems. The pertinent areas were
- a) the primary guidance and navigation section (PGNS) in gudance, navigation and control system (GN&CS)
- b) the descent propulsion section (DPS) in the main propulsion system (MPS)
- c) the reaction control system (RCS)
- d) the landing radar (LR) which did[altitude and ground speed measurements in the RS (Radar Systems)
- e) and the landing point designator (LPD) which was used to landing point targeting through the commander's window in LM
1) PGNS (PRIMARY GUIDANCE AND NAVIGATION SYSTEM)
|LM PGNS parts (IMU - Inertial Measurement Unit, LGC- LM Guidance Computer, ECDU - Electronics Coupling Data Unit, G&N - Guidance & Navigation, PSA - Power Servo Assembly), see also here|
The PGNCS consisted of two major subsystems: an inertial measurement unit (IMU) and a computer. The IMU is the navigation sensor, which incorporates accelerometers and gyros to sense changes in LM velocity and attitude. The IMU sends this information to the computer, which contains preprogramed logic for navigation, for calculation of guidance commands, for sending steering commands (by means of the digital autopilot (DAP)) to the DPS (Descent Propulsion System) and the RCS (Reaction Control System), for processing LR (Landing Radar) measurements of LM range and velocity relative to the lunar surface, and for display of information to the crew. The crew controls the mode of computer operation through a display and keyboard (DSKY) assembly.
2) DPS (DESCENT PROPULSIONS SYSTEM)
The DPS, which contained the rocket engine used for lunar descent and its controls, consists of a throttle and a gimbal drive capable of ±6° of motion. The engine has a maximum thrust of approximately 10 000 pounds (nominal engines varying from 92.5 to 95.5 percent of the design thrust of 10 500 pounds). The maximum thrust level is referred to as the fixed throttle position (FTP) and is used for efficient velocity reduction during the braking phase. The throttle can be controlled automatically by the PGNCS guidance commands or by manual controls. The descent engine is throttle able between 10 and 60 percent of design thrust for controlled operations during the approach and landing phases. The gimbal drive is controlled automatically by the DAP for slow attitude-rate commands. For high-rate changes, the DAP controls the RCS, which consists of four groups of four small control rockets (100 pounds of thrust each) mounted on the LM to control pitch, roll, and yaw.
3) LR (LANDING RADAR)
|LR Antenna Assembly|
The LR, mounted at the bottom rear of the LM, is the navigation sensor which provides ranging and velocity information relative to the lunar surface. The LR consists of four radar beams, one beam to provide ranging measurements and three beams to provide velocity measurements. This beam pattern, which is illustrated relative to the LM body axis system in figures 3(a) and 3(b), can be oriented in one of two positions, as shown in figures 3(c) and 3(d). Position 1 (fig. 3(c)) is used in the braking phase of the descent when the LM is oriented near the horizontal. Position 2 (fig. 3(d)) is used during the approach and landing phases of descent when the LM nears a vertical attitude. The guidance computer converts the ranging information to altitude data and updates its navigated position every 2 seconds. The guidance computer also converts the velocity measurement along each radar beam to platform coordinates and updates a single component of its navigated velocity every 2 seconds; thus, 6 seconds is required for a complete velocity update. The LR data are weighted before they are incorporated into the guidance computer (ref. 7).
4) LPD (LANDING POINT DESIGNATOR)
|Landing Point Designator as seen by the LM commander|
The one system to be described is a grid on the commander's forward window called the LPD. The window is marked on the inner and outer panes to form an aiming device or eye position. During the approach and landing phases, the computer calculates the look angle (relative to the forward body axis ZB) to the landing site and displays it on the DSKY. The commander can then sight along the angle on the LPD (zero being along body axis ZB) to view the landing area to which he is being guided. If the commander desires to change the landing area, he can make incremental changes inplane or cross range by moving the hand controller in the appropriate direction to provide input to the computer. Cross-range position is changed in 2° increments, and inplane position is changed in O. 5° increments.
5) RCS (REACTION CONTROL SYSTEM)
|RCS (Reaction Control System) Components|
|RCS Propellant tank for fuel and oxidizer. Helium around the bladder forced the propellant to exit and no pumps were required.|
LM MAIN PROPULSION SUBSYSTEM /6/ /7/
The Main Propulsion Subsystem (MPS) consists of two separate complete, and independent propulsion sections: the descent propulsion section and the ascent propulsion section.
|LM ascent (upper) and descent (lower) parts. The descent part only contained parts that were needed before ascent from the Moon.|
1) The descent propulsion section provides propulsion for the LM from the time it separates from the CSM (Command Service Module) until it lands on the lunar surface.
2) The ascent propulsion section lifts the ascent stage off the lunar surface and boosts it into orbit.
|Reaction Control System (RCS) schematic|
Each propulsion section consists of a liquid-propellant, pressurefed rocket engine and propellant storage, pressurization, and feed components. For reliability, many vital components in each section are redundant. In both propulsion sections, pressurized helium forces the hypergolic propellants from the tanks to the engine injector. Both engine assemblies have control valves and trim orifices that start and stop a metered propellant flow to the combustion chamber upon command, an injector that determines the spray pattern of the propellants as they enter the combustion chamber, and a combustion chamber, where the propellants meet and ignite. The gases produced by combustion pass through a throat area into the engine nozzle, where they expand at an extremely high velocity before being ejected. The momentum of the exhaust gases produces the reactive force that propels the vehicle.
|Ascent Propulsion System was more simple than the DPS|
|Descent Propulsion System with heat exchangers|
The more complicated tasks (adjustable thrust) required of the descent propulsion section dictate that the descent section be the larger and more sophisticated of the two propulsion sections. In the modified LM vehicle the descent propellant tanks hold 18,697 pounds of usable propellants (compared to 17,489 in the original LM), and amount that is more than three times that of the ascent propulsion section. The descent engine is almost twice as large as the ascent engine, produces more thrust (almost 10,000 pounds at full throttle), is throttleable for thrust control, and is gimbaled for thrust vector control. The ascent engine, which cannot be tilted, delivers a fixed thrust of 3,500 pounds, sufficient to launch the ascent stage from the lunar surface and place it into a predetermined orbit.
|Descent engine functional diagram|
The ascent and descent propulsion sections, as well as the RCS, use identical fuel/oxidizer combinations, In the ascent and descent propulsion sections, the injection ratio of oxidizer to fuel is approximately 1.6 to 1, by weight.
The fuel is a blend of hydrazine (N2H4) and unsymmetrical dimethylhydrazine (UDMH), commercially known as Aerozine 50. The proportions, by weight, are approximately 50 % hydrazine, and 50 % dimethylhydrazine.
The oxidizer is nitrogen tetroxide (N204). It has a minimum purity of 99.5 % and a maximum water content of 0.1 %.
|LM Panel 1 (Commander's side) upper part contains fuel monitoring displays|
PROPELLANT QUANTITY GAGING SYSTEM
The propellant quantity gaging system enables the astronauts to monitor the quantity of propellants remaining in the four descent tanks. It is in operation during the final powered descent phase, from start of the braking maneuver (10 seconds after engine turn-on) until lunar touchdown. The system consists of four quantity-sensing probes with low-level sensors, a control unit, two quantity indicators, a switch that permits the astronauts to select a set of tanks (one fuel and one oxidizer) to be monitored, and a low-level warning light. The low-level sensors provide a discrete signal to cause the warning light to go on when the propellant level in any descent tank is down to 9,4 inches, an amount sufficient for 2 minutes of engine burn at hover thrust (approximately 25%).
|LM Panel 1 is located on the commander's side in the Lunar Module|
Before earth launch, the LM propellant tanks are only partly pressurized (less than 230 psia), so that the tanks will be maintained within a safe pressure level under the temperature changes experienced between filling and launch. At initial engine start, the ullage space in each propellant tank requires additional pressurization. This initial pressurization is accomplished with a relatively small amount of helium stored at ambient temperature and at an intermediate pressure.
|Helium tank (this one from the service module)|
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 10 psi per hour maximum), the initial loading pressure is planned so that the supercritical helium will be maintained within a safe pressure/ time envelope throughout the mission.
The supercritical helium tank has a burst disk assembly and an internal helium/helium heat exchanger. The burst disk assembly 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 rupture 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. In addition to this venting arrangement, the modified LM has a controlled bleed vent in parallel. This bleed vent - a series of stacked perforated plates - permits a small controlled amount of supercritical helium to be bled, extending the supercritical helium tank maximum standby time from 131 hours to 190 hours.
|LM propulsion group 1967|
To open the pressurization path to the propellant tanks, an explosive helium isolation valve and two propellant compatibility values must be fired. The helium isolation valve is automatically fired 1.3 seconds after the descent engine start command is issued. The time delay prevents the supercritical helium from entering the fuel/helium heat exchanger until propellant flow is established so that the fuel cannot freeze in the heat exchanger. The supercritical helium enters the two-pass fuel/helium heat exchanger where it is slightly warmed by the fuel. The helium then flows back into a heat exchanger in the supercritical helium tank where it increases the temperature of the supercritical helium in the tank. Finally, the helium flows through the second loop of the fuel/helium heat exchanger where it is heated to operational temperature.
After flowing through a filter, the helium enters a pressure regulator system which reduces the helium pressure to approximately 245 psi. This system consists of two parallel, redundant regulators. The regulated helium then enters parallel paths, which lead through quadruple check valves into the propellant tanks. The quadruple check valves, consisting of four valves in a series-parallel arrangement, permit flow in one direction only. This protects upstream components against corrosive propellant vapors and prevents hypergolic action due to backflow from the propellant tanks.
A relief valve, which opens at approximately 268 psia, protects each propellant tank against overpressurization. A thrust neutralizer prevents the gas from generating unidirectional thrust, Each relief valve is paralleled by two series-connected vent valves, After landing, the astronauts open the vent valves to relieve pressure buildup in the tanks.
|John Houbolt explains LOR concept|
Premission planning for Apollo lunar module (LM) descent and ascent started in 1962 with the decision to use the lunar orbit rendezvous (LOR) technique for the Apollo lunar-landing mission. The LOR concept advanced by Houbolt and others is defined in Wikipedia. The technique allowed optimization of both the design of LM systems and trajectories for orbital descent to and ascent from the lunar surface.
|Lunar-orbit rendezvous LOR concept|
The LM descent was designed to be accomplished in two powered-flight maneuvers: the descent orbit insertion (DOI) maneuver and the powered-descent maneuver. The DOI maneuver, a short or impulse-type transfer maneuver, is performed to reduce the orbit altitude of the LM from the command and service module (CSM) parking orbit to a lower altitude for efficiency in initiating the longer, more complex powered-descent maneuver.
The basic trajectory design for the powered descent was divided into three operational phases: an initial fuel-optimum phase, a landing-approach transition phase, and a final translation and touchdown phase. The initial trajectory analysis which led to this design was performed by Bennett and Price (ref. 3). In reference 4, Cheatham and Bennett provided a detailed description of the LM descent design strategy. This description illustrates the complex interactions among systems (guidance, navigation, and control; propulsion; and landing radar), crew, trajectory, and operational constraints. As LM systems changed from design concept to hardware, and as operational constraints were modified, it became necessary to modify or reshape the LM descent trajectory; however, the basic three-phase design philosophy was retained. (See previous part of this article for a more detailed description of the powered descent).
DESCENT STAGE /1/
The descent stage was the unmanned portion of the LM (Lunar Module); it represented approximately two-thirds of the weight of the LM at the Earth-launch phase. In addition to containing the descent propulsion section, the descent stage was designed to:
- support the ascent stage
- provide storage to support scientific equipment and the lunar roving vehicle used on the lunar surface
- provide attachment of the landing gear
- serve as the ascent stage launching platform
|Lunar Module (LM) Descent Stage|
The descent stage structure of aluminum-alloy, chemically milled webs provides attachment and support points for securing the LM within the spacecraft - Lunar Module adapter (SLA).
|Spacecraft LM Adapter (SLA) separation parts|
The descent stage structure consists of two pair of parallel beams arranged in a cruciform, with a deck on the upper and lower surfaces approximately 65 inches apart. The ends of the beams are closed off by end closure bulkheads to provide five equally sized compartments: a center compartment, one forward and one aft of the center compartment, and one right and one left of the center compartment. The center compartment houses the descent engine. Descent engine oxidizer tanks are housed in the forward and aft compartments; descent engine fuel tanks, in the side compartments. The entire basic structure is enveloped by a thermal and micrometeoroid shield.
|LM Descent Stage Parts|
Five EPS batteries and two electrical control assemblies (ECA's) are mounted on the rear bulkhead (-Z) of the aft compartment. Four plume deflectors truss mounted on the descent stage divert the plume of the downward-firing RCS thrusters away from the descent stage. A landing gear attenuates landing impact and supports the vehicle. The deflectors in Quadrants I and IV have been shortened to provide the necessary clearances for the payload and MESA, respectively. The supporting trusses have also been modified.
Descent stage Primary Job /15/
The Descent stage's primary job was to support a powered landing and surface extravehicular activity. When the excursion was over, it served as the launch pad for the ascent stage. Octagon-shaped, it was supported by four folding landing gear legs, and contained a throttleable Descent Propulsion System (DPS) engine with four hypergolic propellant tanks.
|LM Landing Radar (LR) under the descent stage|
A continuous-wave Doppler radar antenna was mounted by the engine heat shield on the bottom surface, to send altitude and rate of descent data to the guidance system and pilot display during the landing. Almost all external surfaces, except for the top, platform, ladder, descent engine and heat shield, were covered in amber, dark (reddish) amber, black, silver, and yellow aluminized Kapton foil blankets for thermal insulation.
Kapton is a polyimide film developed by DuPont that remains stable across a wide range of temperatures, from −269 to +400 °C (−452 – 752 °F / 4 – 673 K). Kapton is used in, among other things, flexible printed circuits (flexible electronics) and thermal micrometeoroid garments (the outside layer of space suits).
The descent stage thermal shield combines multiple layers of aluminized mylar and H-film with an outer skin of H-film. In areas where micrometeorite protection is required, one layer of black-painted inconel is used as skin. The shield is mounted on supports, which keep it at least 1/2 inch away from the main structure. The supports have low thermal conductivity. A base heat shield, composed of titanium with a blanket of alternate layers of nickel foil and fiberfax outside, protects the bottom of the descent stage from engine heat, In addition, the engine compartment is protected by a titanium shield with a thermal blanket of multiple layers of nickel foil and fiberfax under an outer blanket of H-film.
|Apollo 16 Lunar Module Foil /16/|
You probably remember the lunar modules, wrapped in bright gold like a present to the cosmos. I always thought that there was one layer of foil to reflect the harsh sunlight in space.
I was surprised to see that this blanket from the Apollo 16 LM was made of 26 layers, of different colors and thicknesses.
As I recently learned, the foil was also a thermal blanket, not just a reflector. Earlier this month, EDN Magazine interviewed Grumman’s Ross Bracco, one of 25 engineers who began development of the LEM, as it was first called:
Still another major challenge Bracco and his team faced was the fact that the LEM was expected to land on the sunny side of the lunar surface, which meant an environmental temperature of 250°F and a shade temperature of -250°F. A low-cost technique was needed to insulate and protect the LEM's structural materials, including the landing feet. The team decided to use 12 to 18 layers of Kapton or aluminized Mylar material sandwiched together in a 70°F earth clean room and trap the air with a special sealing tape. This trapped air remained permanently at 70°F and was used in many areas of the LEM, including the cupped landing feet. The ‘foil’ around much of the LEM was made with 2- and 5-mil aluminized Kapton film.
|The LM "porch"|
|LM Descent Stage - Equipment Locations|
Equipment for the lunar exploration was carried in the Modular Equipment Stowage Assembly (MESA), a drawer mounted on a hinged panel dropping out of the lefthand forward compartment.
Besides the astronaut's surface excavation tools and sample collection boxes, the MESA contained a television camera with a tripod; as the commander opened the MESA by pulling on a lanyard while descending the ladder, the camera was automatically activated to send the first pictures of the astronauts on the surface back to Earth. A United States flag for the astronauts to erect on the surface was carried in a container mounted on the ladder of each landing mission.
The Early Apollo Surface Experiment Package (EASEP) (later the Apollo Lunar Surface Experiment Package (ALSEP)), was carried in the opposite compartment behind the LM.
An external compartment on the right front panel carried a deployable S-band antenna which, when opened looked like an inverted umbrella on a tripod. This was not used on the first landing due to time constraints, and the fact that acceptable communications were being received using the LM's S-band antenna, but was used on Apollo 12 and 14.
|LM S-Band Erectable Antenna Deplaoyment Sequence|
A hand-pulled Modular Equipment Transporter (MET), similar in appearance to a golf cart, was carried on Apollo 13 and 14 to facilitate carrying the tools and samples on extended moonwalks.
|Modular Equipment Transporter (MET)|
On the extended missions (Apollo 15 and later), the antenna and TV camera were mounted on the Lunar Roving Vehicle, which was carried folded up and mounted on an external panel.
|Lunar Roving Vehicle (LRV)|
Compartments also contained replacement Portable Life Support System (PLSS) batteries and extra lithium hydroxide canisters on the extended missions.
DESCENT STAGE DATA
minus landing probes:
|8.59 ft (2.62 m)|
minus landing gear:
|12.83 ft (3.91 m)|
including landing gear:
|31.0 ft (9.4 m)|
|22,783 lb (10,334 kg)|
|Water:||one 151 kg (333 lb) storage tank|
|DPS propellant mass:||18,000 lb (8,200 kg)|
|DPS engine:||TRW LM Descent Engine (LMDE), TRW TR-201|
|DPS thrust:||10,125 lbf (45,040 N), throttleable between 10% and 60% of full thrust|
|DPS propellants:||Aerozine 50 fuel / nitrogen tetroxide oxidizer|
|DPS pressurant:||one 49-pound (22 kg) supercritical helium tank at 1,555 psi (10.72 MPa)|
|DPS specific impulse:||311 s (3,050 N·s/kg)|
|DPS delta-V:||8,100 ft/s (2,500 m/s)|
|Batteries:||four (Apollo 9-14) or
five (Apollo 15-17) 28-32 V,
415 Ah silver-zinc batteries;
135 lb (61 kg) each
The heart of the Lunar Module (LM) descent stage was the Descent Propulsion System (DSP) which was the TRW manufactured Descent Engine and its various control systems. Also the Reaction Control System (RCS) with its jets and controls was important. The RCS jets took mainly care of the continuous attitude control but during burns also the descent engine gimbaling contributed to that. The ascent engine was not gimbaled nor was it adjustable.
DESCENT PROPULSION SYSTEM /5/
Because of the complexity caused by the throttling and gimbaling requirements and the use of an advanced pressurization-system concept, the DPS was one of the last Apollo propulsion systems to reach developmental maturity. Because the DPS was also a pressure-fed earth-storable hypergolic-propellant design, as were the other Apollo spacecraft propulsion systems, DPS development did benefit by some of the knowledge gained in the earlier development of other systems.
Based on the experience with the DPS, system-to-system interface requirements are very critical, and a periodic review of system interface requirements should be conducted. For example, the guidance and navigation system on the LM-1 flight had a AV monitor that terminated the engine burn if a certain AV threshold was not met. This burn termination might have been avoided if the LM-1 engine-start characteristics had been compared with the guidance and navigation program AV monitor.
|The final DPS design configuration of the injector|
After approximately 18 months of parallel engine development by two different contractors using these two throttling concepts, the variable-area-injector throttling concept was selected (January 1965). The final design configuration of this injector is shown in the above figure. The basic reasons for this selection were that the fixed-area injector with helium injection experienced significant problems with combustion instability during development, and reduced-thrust performance was lower than that with the variable-area injector.
|Bipropellant variable area injector operation principle. /18/|
Maximum-operational thrust occurred at a fixed throttle point (FTP) that was calibrated for optimum-mixture ratio and injector pressure difference in settings. The FTP was optimized at 92.5 percent of maximum-rated thrust (10 500 lb). The initial design attempted to provide variable thrust near the maximum-rated thrust; however, excessive problems with mixture-ratio control and throat erosion caused the selection of an FTP maximum-thrust setting, which was compatible with mission requirements.
Thrust between 65 and 92.5 percent was considered a nonoperating region, although the engine can be physically throttled to any point in that region. The minimum-throttle point was 10 percent of the rated thrust.
|LM Descent Engine (TRW)|
The fuel-actuated shutoff valves were parallel-series redundant ball valves to provide fail-safe capability. The fuel and oxidizer injector manifold was sized to achieve an oxidizer lead during the start transient.
The thrust chamber was ablatively cooled to an area ratio of 16: 1. The nozzle extension was a sheet-metal columbium-alloy skirt extending from 16:1 to 47.5:1 area ratio. Cooling of the nozzle extension was by radiation.
The throttle actuator was a triple-redundant, electrically driven device. Accomplished by an electrical-null-balance method, direct-current command signals were compared to throttle-position indication and throttle-position control. A more detailed description of the throttle actuator is provided in reference 1.
Lunar Module "1000 s" TRW Descent Engine /4/
The LMDE is a pressurefed, bipropellant, variable thrust, gimballing, chemical rocket engine with a maximum thrust of 9850 lbs, throttleable down to approximately 1000 lbs. The throttling is obtained by means of dual, variable area, cavitating venturi flow control valves mechanically linked to a variable area injector. Ignition is hypergolic in the combustion chamber since the propellants are nitrogen tetroxide (N204 ) and a 50-50 mixture of hydrazine (N2H4 ) and unsymmetrical dimethylhydrazine (UDMH).
|LM Descent Engine|
The time available for developing and qualifying the LMDE for manned flight was short. To assure that unforeseen problems did not develop during the latter part of the development and qualification phase, it was important to plan an adequate component test program early in the engine's development. This program included tests of details and subassemblies to locate weaknesses in the design. Upon correction of these deficiencies, the next level of assembly was tested. Then, step by step assurance was developed that the complete engine performed as required.
|LM Descent Engine Parts|
Details of the mechanical design of the LMDE are described later in the paper. All aspects of the design, including discussion of the critical environments and the development problems anticipated, are presented along with the component tests used to find the solutions to these problems.
|1000 s TRW Descent Engine|
The total firing time requirement for the LM Descent Engine is 1000 seconds. This is composed of 90 seconds of acceptance testing plus 910 seconds of duty cycle. A nominal duty cycle is shown in Figure 4. Temperature of the titanium combustion chamber case does not exceed 800°F during duty cycle firing. /4/
|Descent Engine Detailed View|
The major percentage of the fuel is injected axially into the combustion chamber through the annular orifice of the variable area concentric injector, while the remainder is distributed to the 36 ports of the barrier cooling ring where it is injected along the wall of the chamber. The oxidizer enters the center of the injector assembly, flows down and is directed into the combustion chamber radially through the variable area ports in a horizontal disc composed of 36 jets which impinge on and ignite with the sheet of fuel.
|LM Descent Engine Functional Diagram|
Thrust control is accomplished by adjusting the propellant flow rates and varying the area of the injection ports to maintain near uniform velocities. Electrical control signals from the Lunar Module are received by the electromechanical throttle actuator which converts the signal to a linear position of the actuator screw jack. Attached to the top of the screw jack is a cross beam. The right side of the cross beam is connected directly to the oxidizer flow control valve through a flexural element. The left side of the cross beam is connected to the fuel flow control valve through the mixture ratio trim linkage, establishing the desired motion of the fuel flow control valve relative to the motion of the oxidizer flow control valve. This produces the required mixture ratio of the propellants.
This linkage permits adjustment of the mixture ratio during acceptance tests of the engine.
Also attached to the cross arm by means of two flexures is a walking beam, pivoted off the manifold assembly, which translates motion from the throttle actuator assembly to the injector assembly. Motion of the movable sleeve in the injector simultaneously changes the gap in the fuel orifice and the size of the ports in the oxidizer outlets.
The manifold assembly which provides the mounting points for the throttle mechanism is the distribution network for the propellants and the end closure of the combustion chamber.
A welded titanium shell which bolts to the manifold forms the structural element of the combustion chamber and the throat of the engine. The phenolic-fiberglass ablative liner provides the thermal barrier between the burning propellants and the structural case. Attachments are provided in the throat area for the gimbal assembly. The chamber and ablative liner extend to the 16:1 area ratio point where a flanged joint provides for attachment of the nozzle extension. The nozzle extension is a radiation-cooled, bell-shaped cone which provides for expansion of the gases from the 16:1 area ratio at the attachment to the combustion chamber to the 47.5:1 area ratio exit.
|Thrust Control Assembly (TRW)|
The linkage between the throttle actuator, the flow control valves and the variable area injector drive mechanism requires high precision. The ratio between the motion of the flow control valves and the injector is 5:1. Total motion of the injector sleeve is 0.15 inch. A ten per cent increment in throttle setting is therefore a motion of only 0.003 inch. In order to maintain the position relationship between the injector sleeve and the flow control valves, looseness in the pivot bearings of the linkage had to be virtually eliminated. Another constraint on the design was that the joints of the mechanism be capable of motion in a complete vacuum.
|LM Descent Engine|
The combustion chamber consists of an ablative-lined titanium alloy case to the 16:1 area ratio. Fabrication of the 6Al-4V alloy titanium case is accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. Thickness of the shell is a uniform 0.035 inch except at the upper end where the head end bolts on, at the weld joint and at the lower flange where the nozzle extension attaches. On either side of the throat a pair of flanges are provided integral with the case for attachment of Z-shaped aluminum rings. These rings form the structural members for attachment of the gimbal trunnions.
|LMDE Sectional View (TRW)|
Since the titanium chamber is one piece, the ablative liner is fabricated in two segments and installed from either end. A metal locking element positively locks both halves of the liner together as they are installed. This lock is redundant since the shape of the nozzle extension is such that the ablative liner is retained in the exit cone during transportation and launch and boost. During engine firing, thrust loads force the exit cone liner against the case.
The landing gear is of the cantilever type; it consists of four assemblies, each connected to an outrigger that extends from the ends of the structural parallel beams. The landing gear assemblies extend from the front, rear, and both sides of the descent stage. Each assembly consists of struts, trusses, a footpad, and lock and deployment mechanisms. The left, right, and aft footpad has a lunar surface sensing probe. A ladder is affixed to the forward gear assembly.
|Lunar Module front leg was equipped with a ladder|
The landing gear attenuates the impact of a lunar landing, prevents tipover, and supports the vehicle during lunar stay and lunar launch. Compression loads are attenuated by a crushable aluminum-honeycomb cartridge in each primary strut. Landing impact is attenuated to load levels that preserves the vehicle structural integrity. At earth launch, the landing gear is in the retracted condition. When the Commander, in the vehicle, operates the landing gear deployment switch, the landing gear uplocks are explosiveIy released and springs in the deployment mechanism extend the landing gear. Once extended, each gear assembly is locked in place by two downlock mechanisms. The lunar surface sensing probe is an electromechanical device, The probes are retained in the stowed position against the primary strut until landing gear deployment. During deployment, mechanical interlocks are released, permitting spring energy to extend the probes below the footpad. At lunar contact, two mechanically actuated switches in each probe energize lights to advise the crew to shut off the descent engine.
PRIMARY GUIDANCE FUNCTIONAL DESCRIPTION
The GN&CS (Guidance Navigation & Cotrol (Sub)System) comprises two functional loops, each of which is an independant guidance and control path. The primary guidance path (PGNS) contains elements necessary to perform all fimctions required to complete the lunar mission. If a failure occurs in this path the abort guidance path (AGS) can be substituted.
|The whole LM control system consisted of 12 subsystems, the GN&CS subsystem consisted of several sections such as PGNS and AGS|
Primary Guidance Path
The primary guidance path comprises the PGNS (Primary Guidance and Navigation Section), CES (Control Electronics Section), LR (Landing Radar), RR (Renezvous Radar), and the selected propulsion section required to perform the desired maneuvers. The CES routes flight control commands from the PGNS and applies them to the descent or ascent engine (MPS), and/or the appropriate thrusters (RCS). The IMU which continuously measures attitude and acceleration, is the primary inertial sensing device of the vehicle. The LR senses slant range and velocity. The RR coherently tracks the CSM to derive LOS range, range rate, and angle rate. The LGC uses AOT star-sighting data to align the IMU. Using inputs from the LR, IMU, RR, TTCA' s, and ACA' s, the LGC solves guidance, navigation, steering, and stabilization equations necessary to initiate on and off commands for the descent and ascent engines, throttle commands and trim commands for the descent engine, and on and off commands for the thrusters.
|ACA Attitude Control Assembly (Rotation Control with a Joystick)|
Control of the vehicle when using the primary guidance path, ranges from fully automatic to manual. The primary guidance path operates in the automatic mode or the attitude hold mode. In the automatic mode, all navigation, guidance, stabilization, and control functions are controlled by the LGC. When the attitude hold mode is selected, the astronaut uses his ACA to bring the vehicle to a desired attitude. When the ACA is moved out of the detent position, proportional attitude-rate or minimum impulse commands are routed to the LGC. The LGC then calculates steering information and generates thruster commands that correspond to the mode of operation selected via the DSKY. These commands are applied to the primary preamplifiers in the ATCA, which routes the commands to the proper thruster. When the astronaut releases the ACA, the LGC generates commands to hold this attitude. If the astronaut commands four-jet direct operation of the ACA by going to the hardover position, the ACA applies the command directly to the secondary solenoids of the corresponding thruster.
|LM GN&CS subsystem handled guidance, navigation and control|
In the automatic mode, the LGC (LM Guidance Computer) generates descent engine throttling commands, which are routed to the descent engine via the DECA (Descent Engine Control Assembly). The astronaut can manually control descent engine throttling with his TTCA (Translation and Thrust Control Assembly). The DECA sums the TTCA throttle commands with the LGC throttle commands and applies the resultant signal to the descent engine. The DECA also applies trim commands, generated by the LGC, to the GDA' s (Gimbal Drive Actuators) to provide trim control of the descent engine. The LGC supplies on and off commands for the ascent and descent engines to the S&C control assemblies. The S&C control assemblies route the ascent engine on and off commands directly to the ascent engine, and the descent engine on and off commands to the descent engine via the DECA.
In the automatic mode, the LGC generates +X-axis translation commands to provide ullage. In the manual mode, manual translation commands are generated by the astronaut, using his TTCA. These commands are routed, through the LGC, to the ATCA and on to the proper thruster.
IMU (Inertial Measurement Unit) /10/
The design of the inertial subsystem required for the navigation and guidance of the Apollo spacecraft was a responsibility separate from spacecraft vehicle design. Early mission-error analysis indicated that accelerometers and gyros of the Polaris Mark II system had performance characteristics adequate for the Apollo lunar mission. The Apollo inertial system thus evolved from basic Polaris Mark II designs. This decision was heavily based on the initial requirement for an Earth-orbital flight in late 1963.
|Apollo CMS and LM IMU mechanical gyros (there were no electronic gyros available at that time)|
The Block II and LM inertial subsystems consisted of the IMU, the electronic coupling data unit (ECDU), the PSA (Power Servo Assembly), a navigation base, the pulsed integrating pendulous accelerometer (PIPA) electronics assembly (PEA) in the CM, and the pulsetorquing assembly (PTA) in the LM.
The inertial subsystem performs three major functions:
1) measures changes in spacecraft attitude,
2) assists in generating steering commands, and
3) measures spacecraft velocity changes caused by thrust or atmospheric drag.
To accomplish these functions, the IMU provides an inertial reference consisting of a stable member having three degrees of freedom that is stabilized by three integrating gyros. When the inertial system is operated before launch, the stable member is alined through a gyrocompassing routine; during flight, the stable member is alined by sighting the optical instruments on stars. If the inertial subsystem is operated for several hours, realinement may be necessary because the gyros that maintain the spacereferenced stable member may drift and cause an error in flight calculations.
|The navigation base was located above the astronauts in LM|
The design decisions concerning the inertial subsystems were heavily influenced by the plan (late 1961) to fly in 2 to 3 years. That period of time would not permit a complete new inertial system development. Thus, the design of the Block I inertial system was based on the Polaris Mark II system. Both the gyro and accelerometer used basic Polaris designs with minor mechanical and electrical changes. The early programmatic decisions also committed the Apollo inertial program to the competence and experience of the Polaris Mark II institutional and industrial team. The new areas of development were for the Apollo manned-mission design requirements of (1) inflight optical alinement interface, (2) pilot moding interface, (3) general-purpose digital-computer gimbal-angle interface, (4) in-flight repair, and (5) packaging and interconnect wiring.
It appears that the schedule push early in the Apollo Program (i. e., fly by late 1963) may very well have led to inadequate systems review before designs were frozen. Within the inertial subsystem, however, the conservative design decisions in most cases were very close to optimum. Use of the Polaris technology certainly reduced the risk and effort required for most elements of the inertial subsystem design. Unrealistic workloads required of the crewmen in maintenance, fault isolation, mode selections, and operational techniques were the primary mistakes corrected in the change from Block I to Block II configurations.
LM Digital Control ("ags and pings")
There were 2 computers in LM to perform control operations. Both were independent from each other and could handle their tasks without the other. These computers were the primary computer (PGNS) called LGC, Lunar Module Guidance Computer (or AGC, Apollo Guidance Computer) which was exactly the same computer that was also in the CSM; and the AGS (Abort Guidance System) computer which was in the abort electronics assembly; this computer could only abort the landing mission (not continue the landing)
|The primary digital computer (LGC or AGC) in LM with DSKY panel|
The PGNS (or DAP) computer LGC was interfaced via the DSKY panel on the commander's side and the AGS computer was interfaced via the panel 6 (DEDA panel) on the LM pilot's side.
|AGS computer main components DEDA, AEA and ASA|
LM Stability and Control /11/
The lunar module stabilization and control, which is part of the lunar module guidance, navigation, and control system, is designed to control vehicle attitude and translation about or along all axes during a lunar module mission. Vehicle attitude and small translations are controlled by selectively firing the 16 reaction control system jets mounted on the ascent stage. The design concepts used were representative of the state of the art, and six different assemblies constituted the subsystem:
1) ATCA, attitude and translation control assembly,
2) DECA, descent engine control assembly,
3) RGA, rate gyro assembly,
4) ACA, attitude controller assembly,
5) TTCA, thrust and translation controller assembly, and
6) GDA, gimbal drive actuator.
|Electrical parts placement in the LM|
The major features of both the PGNCS and AGS modes are summarized in the following brief functional description of the S&C. The LM S&C is part of the LM guidance, navigation, and control system and is designed to control vehicle attitude and translation about or along all axes during an LM mission. Vehicle attitude and small translations are controlled by firing one or more of the 16 RCS jets mounted on the ascent stage. Major translations are accomplished by means of the ascent or descent propulsion engine. Rotation around the LM X-axis, Y-axis, and Z-axis is termed yaw, pitch, and roll, respectively. Movement along an axis is termed translation. The LM axes form a right-handed orthogonal triad.
The LM guidance, navigation, and control system contains two independent guidance sections. The PGNCS is designed to control the LM during all mission phases. The AGS is designed to guide the LM to the command and service module if the PGNCS fails.
|RCS jets worked with or without the main propulsion to hold the attitude|
Primary Mode (PGNS)
When operating in the PGNCS mode, the S&C performs the following functions.
1. Converts RCS jet commands issued by the LM guidance computer (LGC) to the electrical power required to operate the RCS-jet solenoid valves for attitude and translation control
2. Accepts discrete (on-off) descent-engine gimbal commands from the LGC (Upon receipt of an "on" command, the descent engine is gimbaled about its axes at a constant angular rate until the command is removed. )
3. Accepts LGC automatic and manual engine on-off commands and routes them to the propulsion system to start or stop the descent or ascent engine
4. Accepts LGC automatic thrust commands and thrust and translation controller assembly (TTCA) manual thrust commands to control the thrust of the descent engine
5. Provides manual attitude and translational commands to the LGC
|LM DAP Associated Interfaces (DAP and/or PGNS were programs in the LGC)|
The LM digital autopilot (DAP) has interfaces with 16 RCS solenoid driver preamplifiers in the attitude and translation control assembly (ATCA) and with the pitch- and roll-trim gimbal servomotor drives in the descent-engine control assembly (DECA). During the descent-engine burn, the DAP tries to control the spacecraft attitude with the trim gimbal servomotor drive to save RCS propellants; however, if the rather slow trim gimbal servomotor drive does not keep the attitude error within specified bounds, the RCS jets are used until the error is brought within the attitude dead band, and control is then returned to the trim gimbal servomotor drive.
The DAP works in conjunction either with a PGNCS guidance loop to provide an automatic G&C system or with the TTCA and the attitude controller assembly (ACA) for manual G&C. In the latter mode of operation, the DAP provides the crewman with an integrated stabilization and control system for performing translational maneuvers and for maintaining rate command/attitude hold.
The DAP monitors the eight RCS thruster-off signals and chooses the best set of jets to use under the combined conditions of rotational commands, translational commands, and disabled jets. Propellant economy, minimization of the number of RCS jet firings, and operation with detected and undetected jet failures are of primary concern. The LGC will issue automatic engine on, engine off, descent engine thrust magnitude, and descent-engine gimbal position commands, all automatically under program control.
The rate-command/attitude-hold mode is the normal means of astronaut control of the spacecraft. The maximum maneuver rate about any axis of the spacecraft is 20 deg/sec. The ACA acts as an analog device during rate command by producing a voltage proportional to stick deflection. The voltage, which represents commanded rate, is converted to a binary number (quantized at approximately 0.625 deg/sec/bit) and presented to the DAP.
When the stick is out of detent, the DAP tries to match the vehicle rate to the rate commanded by the hand controller. When the rate error is less than the rate dead band (0.4 deg/sec during descent, 1.0 deg/sec during ascent), the jets are no longer commanded on.
However, if the rate error exceeds R specified hound (2,O deg/sec during ascent 1.4 deg/sec during descent), four jets are used to torque the spacecraft. When the ACA is returned to the detent position, the DAP computes the time of jet firing required to zero the rates of the vehicle.
When the rate of the vehicle is brought inside a dead band about zero rate, the contents of the coupling data unit (CDU) registers are transferred to the desired CDU registers, and attitude steering about the newly attained position of the vehicle is begun. If the spacecraft has large pitch- and roll-rate errors, diagonal jets are used; the jets are thus selected efficiently.
The crewman uses the minimum-impulse mode to control the spacecraft with very small rate maneuvers. Each discrete deflection of the ACA 2.5' or more out of detent will cause the DAP to issue commands to the appropriate jets for a minimum impulse. In this mode, an astronaut must anticipate on his own and perform rate damping and attitude steering.
Selection of the minimum-impulse mode enables the crewman to perform an economical low-rate maneuver to a new orientation of the spacecraft. After completing the maneuver, the crewman codes the display and keyboard, and the DAP returns to attitude hold and causes the vehicle to limit cycle, with the normal attitude steering about the new orientation.
Abort Mode (AGS)
When operating in the abort guidance mode, the AGS performs the following functions.
|AGS main components ASA, DEDA and AEA (the computer)|
1. Accepts attitude-error signals from the AGS or manual attitude-rate commands from the ACA and fires the proper RCS jets to achieve attitude control (Rate damping is achieved by summing the error signal or the rate command with the rate gyro feedback signals.)
2. Accepts manual translational commands from the TTCA and fires RCS jets to accelerate the LM in the desired direction'
3. Automatically gimbals the descent engine for trim control
4. Accepts AGS automatic and manual engine on-off commands and routes them to the descent or ascent engine
5. Accepts TTCA manual-throttle commands to control the thrust of the descent engine
The abort system provides the same basic autopilot modes provided by the DAP; that is automatic, rate command/attitude hold, and pulse. In these modes, the ATCA (rather than the LGC) receives signals from the TTCA the ACA, and the AGS (steering errors or attitude reference) and provides the necessary jet-select logic for jet firing. In the pulse mode, a series of minimum-impulse jet firings is provided when the ACA is deflected 2.5 O. This series of pulses contrasts with the single pulse used for this mode in the primary system.
The major elements of the SCS subsystem functional and performance requirements were developed from 1963 to 1965. Periodic meetings were held at the NASA Lyndon B. Johnson Space Center (JSC) (formerly the Manned Spacecraft Center (MSC)) during the requirements definition phase. Interested personnel from MSC NASA Headquarters and three contractors participated in the meetings which were helpful in bringing different viewpoints into focus regarding the establishment of detailed requirements concurrently with hardware design. Before the integrated G&C concept was adopted, the more significant results of these meetings were recorded in the proceedings of the LM SCS meetings for the period from 1963 to early 1964. The implementation of the integrated concept began in September 1964 and is documented in the LM G&C implementation meeting minutes.
ATCA (Attitude and Translation Control Assembly)
|ATCA - Attitude and Translation Control Assembly in LM|
The Stab/Cont ATCA circuit breaker on panel 11, when closed, applies DC power through the Guidance Control-PGNS switch. It also applies DC power through the PGNS Mode Control switch to the ATCA (Attitude and Translation Control Assembly) primary preamplifiers to be used by the PGNS for attitude control. - During the RCS cold fire tests, both ACAs and TTCAs are used to check electrical connections from each controller to the RCS jet drivers in the ATCA (Attitude and Translation Control Assembly).
Apollo 11 communication 1969: MET 112:55:05 McCandless: "Tranquility Base, this is Houston. In the flight plan configuration, we show that the stability-control circuit breaker ATCA (Attitude and Translation Control Assembly) on panel 16 should be open at this time. Over. (Pause)"
|ATCA functional diagram|
The ATCA is composed of four output subassemblies, three analog subassemblies, one power supply, one wiring subassembly, the assembly chassis and cover, electrical connectors, and one elapsed-time indicator. The complete assembly was specified to have a maximum weight of 12 kilograms (27 pounds). The ATCA was mounted on cooling rails in the aft equipment bay of the LM ascent stage.
The ATCA assembly-level developmental tests were performed on two units. Design-feasibility tests were performed on breadboard serial number (S/N) 001s, and design-verification tests were performed on preproduction unit S/N 004, part number (P/N) LSC-300-140-5. During design-verification testing, one failure occurred that required redesign. This failure consisted of a radiofrequency interference at 120 megahertz and its harmonics. The problem was resolved by adding decoupling capacitors at the pulse-ratio modulator. One additional failure was resolved by changing the specification with regard to the distortion of the 800-hertz power-supply voltages that occurred at 150 hertz during audio-conducted susceptibility testing.
Two production assemblies (P/N LSC-300-140-7, S/N 011 and 013) were used in the qualification test program. Endurance qualification tests were performed on unit S/N 011, and design-limit tests were performed on unit S/N 013. The test requirements are contained in CTR LCQ-300-007, dated November 30, 1966, and revision A, dated August 24, 1967.
No significant problems were encountered during qualification testing. However, a postqualification inspection of unit S/N 011 revealed the presence of cracks in some of the solder joints in the interconnection boards used to electrically and physically interconnect a number of cordwood modules, The modules contain copper bus wires that are soldered to the top of the interconnection board; the resulting space between the cordwood and the board is filled with urethane. The cracked joints were caused by mechanical stress resulting from temperature expansion of the urethane filler. The solution to this problem was to strengthen the joint by making it a reflowed convex joint. Because of schedule and cost considerations, a design change to provide stress relief was not made. Instead, all flight attitude and translation control assemblies were modified.
An ATCA design change became necessary after the qualification had been completed. The change was made to eliminate a high-rate-limit-cycle (nonminimum impulse) condition for lightweight ascent conditions. The high-rate limit cycle was caused by marginal control-loop dynamics resulting from a slight change in the RCS thruster characteristics. The ATCA change consisted of altering the pulseratio modulator nonlinearity parameter from a value of 0.1 to 0.3. This change was qualified on ATCA unit S/N 013, after modification, to the requirements of CTR LCQ-300-029, dated July 10, 1968. The ATCA change was effective for LM-4 and subsequent lunar modules.
DECA - Descent Engine Control Assembly /7/ /1/
|DECA - Descent-Engine Control Assemply in LM|
/Apollo 10 Flight Journal:/
[Closing the Stab/Cont, DECA Power circuit breaker on panel 11, provides power to the descent engine control assembly.]
[The DESC ENG CMD OVRD switch on panel 3 is used to keep the DPS engine shut off valves open if DECA power fails. In the Off position, power is removed from these valves.]
[The Stab Cont DECA Pwr circuit breaker on panel 16 is opened to remove DC power from the DECA which controls the descent engine throttle and gimbals.]
[The DECA (Descent Engine Control Assembly) PWR circuit breaker on the Stab/Cont panel 11 on the CDR's side of the LM cabin is closed to provide power to control the descent engine throttle and gimbals.]
|DECA functional diagram|
The DECA processes engine-throttling commands from the astronauts (manual control) and the LGC (automatic control), gimbal commands for thrust vector control, preignition (arming) commands, and on and off commands to control descent engine operation. The DECA accepts engine-on and engine-off commands from the S&C control assemblies, throttle commands from the LGC and the TTCA, and trim commands from the LGC or the ATCA. Demodulators, comparators, and relay logic circuits convert these inputs to the required descent engine commands. The DECA applies throttle and engine control commands to the descent engine and routes trim commands to the gimbal drive actuators.
|Engineer Owen J. Ingram inspects Lunar Module DECA 1969.|
The DECA contains approximately 1000 parts; 95 percent of these parts are packaged as functional entities in 27 cordwood assemblies, and the remaining parts are mounted directly to the chassis. Functionally, the DECA is made up of two trim-error subassemblies, two trim-error malfunction- detection subassemblies, one automatic-throttle subassembly ! one manual-throttle subassembly, one power-switching subassembly, one elapsed-time indicator, two electrical connectors, and the chassis, However, these subassembly functions are not packaged as separate, replaceable, or interchangeable subassemblies. The complete assembly weight was specified to be no greater than 3.33 kilograms (7.35 pounds), and the assembly was mounted on the LM descent stage. No cold- plate cooling was required.
The DECA assembly-level developmental tests were performed on two units. Design-feasibility tests were performed on breadboard model S/N 001, and design- verification tests were performed on preproduction model S/N 004. No significant problems discovered in design-verification tests required redesign. Two production assemblies were used in the DECA qualification test program. Unit S/N 009 (P/N LSC-300-130-5) was used for endurance testing, and unit S/N 012 (P/N LSC-300-130-9) was used for design-limit testing. The part-number differences resulted from a design change made to the trim-fail circuits during the LM critical-design review before the beginning of qualification testing. In the S/N 009 configuration, an automatic trim shutdown was provided for use if the fail circuits detected a failure. In the S/N 012 configuration, a manual disable capability was provided instead of the automatic shutdown. All flight vehicles following LM-1 were equipped with a DECA having the manual disable capability. The qualification test requirements are stated in CTR LCQ-300-006, dated April 23, 1966. No problems encountered during qualification testing necessitated a design change.
A delta-qualification test was performed on retrofitted DECA unit S/N 012 after the previously mentioned DECA testing had been completed. The delta- qualification test was performed to qualify two DECA design changes that had been made because of a requirement for vibration of higher level than that specified by CTR LCQ-300-006. The two design changes were made as a result of problems encountered during vehicle-level tests on LM-1. One change was the addition of a coincident-pulse-detector circuit in the automatic-throttle counter circuit. This change was made to prevent complementing of the counter output caused by simultaneous receipt of thrust-increase and thrust-decrease pulses from the LGC. The other change was made to lower the 4.3-volt direct-current power-fail-monitor threshold. This change was made so that voltage drops in the vehicle cabling that supplies 4.3 volts to the DECA would not result in erroneous operation of the powerfail monitor. The vibration levels were increased as a result of LTR-3 tests, which showed the DECA vibration levels to be higher than those used in the original qualification testing.
RGA (Rate Gyro Assembly)
|RGA - Rate Gyro Assembly in LM|
The RGA consists of three subminiature rate gyros (located so that their sensitive axes form an orthogonal triad), a support and insulator assembly, a component board and connect assembly, and an elapsedtime indicator. The gyro is a conventional aircraft-type rate gyro modified to accept 800-hertz power for compatibility with the primary guidance power frequency. The maximum specified weight of the assembly was 0.9 kilogram (2 pounds). The RGA developmental tests were performed on six assembly-level hardware items. Feasibility tests were performed on two of three units in addition to vibration testing of a mass model and materials and components testing. Design-verification tests and critical-environment tests were performed on four assemblies. No major design deficiencies were encountered during these tests.
Two production assemblies subjected to qualification testing successfully fulfilled all requirements; no major problems were encountered. The test requirements are defined in CTR LCQ-300-008, dated April 25, 1966. After test completion, vibration data from LTA-3 tests revealed higher vibration levels than those used in qualification. Revision A of CTR LCQ-300-008, dated May 25, 1967, reflects these increased levels. Revised vibration tests at these higher levels were performed successfully on one of the RGA qualification units.
An RGA problem that occurred on the Apollo 10 mission (LM-4) apparently was caused by static friction (stiction). The yaw rate gyro "hung up" for approxi mately 40 seconds. Subsequent analysis indicated that contamination was the most likely cause of the problem and that subjection of the gyro to a questionable rework process during manufacture probably contributed to its contamination. Because the manufacture of all gyros had been completed, a stiction test was instituted to screen gyros with potential sticking problems. Another RGA problem, discovered during the LM-6 checkout at the NASA John F. Kennedy Space Center (KSC), was evidenced by a lack of gyromotor synchronization. Although the LM-6 unit had operated properly during previous RGA acceptance and vehicle tests, an analysis of the unit revealed a deficient gyromotor hysteresis ring. A low-voltage-margin test was instituted to screen flight units for this deficiency.
The low-voltage-margin test seems to be an effective screen for marginal motor characteristics. However, the stiction test is not believed to be effective in detecting all contaminated units. A stripdown and particle count performed on three gyros revealed counts varying from approximately 400 to 1300 and sizes ranging from 0.025 to 0.889 millimeter (0.001 to 0.035 inch). Previous stiction tests on these gyros indicated a slight stiction on two and no stiction on the third. Because the stiction screening test was not effective in detecting all contaminated units, a plan was developed to reduce gyro contamination levels. This plan required that the gyros be returned to the vendor for rework. In rework, the gyros were disassembled and internal cavities and dead spaces sealed off to enhance gyro cleaning. The gyros then were flushed to specified contamination levels and refilled under white-room conditions with filtered fluid. The stiction screening test was continued following the cleaning rework. No further stiction problems were encountered either in screening or in flight
ACA (Attitude Controller Assembly)
|ACA - Attitude Controller Assembly used in LM and CSM|
The principal components of the ACA are three linear- output transducers, one three-axis rotational control mechanism, a housing and grip, two electrical connectors, and 43 switches. The specified maximum permissible weight, including connectors and cables, was 2.15 kilograms (4.75 pounds). Four attitude controller assemblies were used in the development-test program, which consisted of two phases. Phase I tests were performed on two preprototype assemblies fabricated in the model shop. The primary objective of Phase I testing was to obtain enough test data to establish an acceptable prototype design. The Phase I1 development tests were performed on a prototype controller; a second unit was available as a backup. The primary objective of these tests was to provide hardware-design verification in preparation for a complete mission simulation. The developmental testing did not reveal any major design problems. However, some test discrepancies required design or procedural changes (or both). For example, one procedural change required that the switch packages be subjected to high temperature (339 K (150O F)) during ACA buildup to ensure correct switch overtravel adjustment. This particular change was not effective in ensuring correct overtravel adjustment because a high percentage of attitude controller assemblies exhibited improper switch adjustment when thermal-vacuum acceptance testing was later imposed as an acceptance requirement. This problem is discussed in the following paragraphs.
Two production assemblies were subjected to the qualification tests, as required by CTR LCQ-300-001, revision B, dated November 22, 1966. Both units successfully completed testing without any significant discrepancies. A problem was encountered at low temperature (255 K (Oo F)) with the pitch-axis torque hysteresis, which was below the required 60-percent minimum value. It was determined that at low temperature, the minimum hysteresis level in the pitch axis could be changed to 50 percent without serious effect on ACA use. Therefore, the requirement was changed, and a hardware-design change was not necessary.
A switch-actuation problem discovered during LTA-8 thermal-vacuum tests at MSC was caused by improper switch adjustment, or turn-in, during ACA final assembly and calibration; improper adjustment prevented switch actuation under thermal or vacuum conditions (or both). This problem was discovered just as the thermal-vacuum acceptance test requirements of contract-change authorization (CCA) 1108 were being imposed on ACA acceptance. Because the manufacture of all assemblies had been completed at the time of CCA 1108 implementation, it was necessary to return the controllers to the vendor for reacceptance. Many of the contrcllers examined contained improperly adjusted switches that did not actuate during thermal and thermal-vacuum conditions. This problem was resolved by reworking all flight attitude controller assemblies to newly developed switch- calibration procedures and by implementing more effective quality inspections. The recalibrated controllers were used on LM-3 and subsequent lunar modules. Much concern was evident in the program regarding the endurance of the ACA centering spring. This concern was stimulated by the switch-adjustment problem and by a spring failure at the NASA George C. Marshall Space Flight Center (MSFC) on one of the development-model assemblies (with uncontrolled usage) that had been provided to MSFC for laboratory use. The spring failed after an estimated 100 000 actuations. It was estimated that a flight ACA would undergo fewer than 1000 ground and flight actuations. However, because of the difficulty in predicting the actual number of actuations and the actual spring life, a redesigned spring was incorporated when the supply of the existing spares was exhausted in spring-life tests. The redesigned spring was installed only in an ACA that required rework at the vendor; other assemblies used springs of the original design.
TTCA (Thrust and Translation Controller Assembly)
|TTCA - Thrust and Translation Controller Assembly used in LM and CSM|
The TTCA was fabricated in-house by the LM contractor. The TTCA contains a position transducer for commanding proportional throttle signals and 12 switches for providing three-axis translational signals to either the primary or the abort control system. A mode- selection lever is provided for selecting either the throttle or the jet mode. The maximum specified TTCA weight was 2.38 kilograms (5.25 pounds).
The TTCA developmental testing was performed on three controllers. Feasibility tests were performed on one controller, and verification tests were performed on two controllers. No major problems were encountered.
Two production assemblies were subjected to the qualification tests required by CTR LCQ-300-003, revision A, dated November 9, 1966. A total of 13 failures occurred during qualification testing; one design change and two procedural changes resulted. The design change resulted from an excessive force spike (more than 66 I 7 newtons (15 pounds), compared with 31 newtons (7 pounds) as the maximum permissible) during X-axis cycling in the throttle-mode portion of the integration and checkout test. The high force occurred after 258 cycles; at that time, the soft- stop parts were replaced. After approximately 250 cycles, the high-force spike occurred again. A failure analysis revealed that the high-force spikes had been caused by a wearing action on the chrome surface of a throttle linkage cam. The spike bungee and the cam were redesigned to reduce cam surface stresses. The redesigned model was subjected to 1580 cycles without degradation or visible wear. This problem has not reappeared on the redesigned units.
A problem occurred during an operational check at ambient conditions when a switch in the endurance-test unit failed to actuate. Results of a subsequent failure analysis showed that the discrepant switch had a release force of zero. Other tests revealed that a maximum reduction of 0.56 newton (2 ounces) in release force could be expected after 2500 to 15 000 cycles. Before this occurrence, the minimum acceptable switch-release force had been specified as 0.56 newton (2 ounces). This problem was resolved by increasing the minimum switch-release force to 1.11 newtons (4 ounces). Another switch-actuation problem was that the switch turn-in was insufficient to accommodate thermal-vacuum conditions. The existing requirement for a turn-in of 30' was not adequate for all environmental conditions and was increased to 45'. The additional turn-in resolved this problem. A tolerance study for the 45' switch adjustment was conducted by the LM contractor.
GDA (Gimbal Drive Actuator)
|GDS- Gimbal Drive Actuator, 2 of these changed the descent engine's gimbal roll and pitch angles|
A qualification test performed on the GDA prototype model demonstrated that the prototype design satisfied the requirements under environmental conditions imposed in the order of occurrence of the applicable environment during a mission. An acceptance test was performed on each assembly to demonstrate that the assembly satisfied all the requirements of the applicable specification and that it conformed to the applicable approved design.
After the qualification test, the GDA was subjected to a reacceptance vibration test (CCA 726) and a thermal-vacuum test (CCA 1108). Operational failure of one unit during thermal-vacuum testing was attributed to low temperature, not to vacuum. Just before that failure, a GDA failure under ambient pressure and temperature at the LM contractor facility was caused by a reduced airgap between the motor- brake polarizer and the brake armature. As a result of testing (CCA 11081, it was determined that many of the units deviated from the performance requirements. The malfunction of the actuator was attributed principally to the motor. The predominant malfunction was coasting of units caused by failure of the brake to engage in the alternating-current motor. The failure rate was approximately 50 percent on coast; however, one unit drew excess power. Although coasting of the actuator did not constitute a problem within the actuator prolonged coasting caused the DECA to indicate a descent-engine trim failure.
The andyses of the Apollo 9 and Apollo 10 missions for a nominal and a faulty GDA brake showedxegligible effects. Therefore it was concluded that a GDA brake failure would not preclude the completion of these missions. However the acceptability of a false caution-and-warning GDA failure indication resulting from a GDA brake failure was still a matter for consideration.
The vendor was directed to modify the GDA. The redesign affected only the motor area and embodied a constant-drag principle. The redesigned actuator was subjected to a delta-qualification test and an extended-life test. When these tests had been successfully completed it was determined that the new design would be flown on Apollo 11 and subsequent missions.
After the new actuators had been installed in LM-5 (Apollo 111, a failure occurred - the actuator did not respond to a start command. The failure resulted from a phase-angle difference between the power circuit of the DECA and the GDA motor. A capacitor assembly was added at the DECA/GDA interface to improve the phase-angle relationship.
A review of the design-verification testy the qualification test the vehicle test, and the mission data supports the fact that the GDA fulfilled all design requirements. The gimbal drive actuators performed all required functions throughout the Apollo 9, 10 and 11 missions. The caution-and-warning light indicating gimbaltrim malfunction was observed on the Apollo 10 mission. The actuation of this caution-and-warning signal was attributed to brake coasting because the GDA performed normally during subsequent use on the Apollo 10 mission.
LM DAP (Digital Auto Pilot)/14/
Several structural-bending problems associated with the LM/CSM RCS were identified. The original design was unstable in bending because of (1) state-estimator time lags and (2) jet-firing logic that sometimes inhibited jet commands for whole sampling cycles. This dynamics problem was identified late in the design because simulation techniques were used exclusively to verify the stability characteristics, and an unfortunate simulator error was made in the representation of the bending dynamics. These problems resulted in erroneous testing results. This design and verification problem illustrates the necessity of performing stability analysis in addition to detailed simulation testing. Acceptable bending stability was obtained by decreasing the state- estimator gains and by increasing the attitude dead band.
Subsystem Verification Tests (SCS)
Because the subsystem has been procured at the black-box or assembly level none of the subsystem-level verification work was done by the equipment vendors. All subsystem-level testing was accomplished in vehicles and in the LM contractor
Full Missicn Engineering Simulator / Flight Control Integration (FMES /FCI) Laboratory.
The FMES/FCI Laboratory was used to perform preinstallation tests on all flight equipment before it was installed in a vehicle. These tests were performed at ambient environmental conditions at the assembly level only. The FMES/FCI Laboratory was used earlier in the Apollo Program for developing preinstallation test procedures and for performing assembly-integration tests. The integration tests, which included the first electrical interfacing of individual assemblies, were performed on early breadboard or preproduction equipment to verify interface compatibility and to help establish subsystem-level test tolerances. No major problems were encountered during these tests.
Closed-loop simulation tests were also performed to verify system dynamic characteristics and to supplement analytical evaluations. The previously noted ATCA change (associated with the high-rate limit cycle) was evaluated in a closed-loop simulation and supported an associated analytical assessment.
The LTA-1 vehicle tests
The LTA-1 vehicle was used for early systems- level testing. In general, the subsystem hardware used was of a preproduction configuration. The testing was discontinued at an early date; no significant hardware or interface problems were discovered during the limited test activities. The tests were useful in developing the vehicle operational checkout procedures.
The LTA-8 thermal-vacuum tests
The LTA-8 thermal-vacuum tests were performed in the thermal-vacuum facility at MSC. The only problem encountered in the stabilization and control system was associated with the ACA switch-actuation deficiency. This problem was resolved as noted in the subsection entitled "Attitude controller assembly.
The LM-2 drop test
The LM-2 drop-test program was conducted at MSC to evaluate subsystem integrity after vehicle drops representing worst-case lunar- touchdown conditions. Individual assemblies had been designed to survive landing shocks safely, and this capability had been demonstrated during the assembly- level qualification testing. However, cabling integrity had not been demonstrated, and these tests served to verify that no problems existed in this area of vehicle design.
MISSION EXPERIENCE (SCS)
All detailed test objectives had been satisfactory accomplished before the Apollo 11 (LM-5) mission, These objectives, pertaining in whole or in part to the SCS, are given in table I. The problems that occurred during each mission are described in the following paragraphs.
The LM-1 Mission (Apollo 5)
The LM-1 mission was unmanned and was flown without an operating AGS. Descent-engine burns, ascent-engine burns, and fire-in-the-hole staging were accomplished by using the SCS in conjunction with the LM mission programer. Attitude control was accomplished with only the rate-stabilization loop because the AGS normally provides the attitude reference. The main-propulsion burns and staging were performed with the SCS because of an LGC software/propulsion interface incompatibility.
|LM-2 is currently on display at the National Air and Space Museum|
The LM-2 Mission
Was never flown. LM-2 is currently on display at the National Air and Space Museum. LM-2 was trop tested so it might have had permanent structural failures due to that.
The LM-3 Mission (Apollo 9)
The LM-3 mission (the first manned LM mission and an Earth-orbital mission) was accomplished with only one SCS discrepancy: a failure indication in the descent trim system. This indication was not an unexpected flight occurrence because the GDA coasting problem, which produces such an erroneous indication, had been experienced with LM-3 during checkout at KSC. (See subsection entitled "Gimbal drive actuator .") An evaluation of this problem in terms of mission effects had been made before the mission, and a decision was made not to replace the actuators with units having the coasting modification. Therefore, all concerned personnel were prepared for this occurrence, and no detrimental mission effects resulted.
|The Rate Gyro was located near the IMU on the Navigation Base.|
The LM-4 Mission (Apollo 10)
The rate gyro "hangup" problem (previously noted in the subsection entitled "Rate gyro assembly") occurred in lunar orbit of the LM-4 mission during descent staging. The total time of abnormal operation was approximately 40 seconds. This problem was identified during data analysis after the mission. Except for this discrepancy, the subsystem performed normally.
The LM-5 Mission (Apollo 11)
The first lunar landing mission (LM-5) was accomplished without any known subsystem discrepancy or problem
CONCLUDING REMARKS (SCS)
Special care should be devoted to mechanical stresses that a given packaging design may place on solder joints. The solder-crack problems experienced on the attitude and translation control assembly, on other lunar module subsystems, and at the NASA George C. Marshall Space Flight Center indicate the need for such design care before a production commitment is made.
Because the mechanical calibration of some equipment may be affected by environmental factors, such equipment should be verified for acceptance by appropriate environmental tests. The problems experienced with attitude controller assembly switch calibration attest to the value of such tests. The gyro contamination problem and its resolution indicate the need for special care in the design and fabrication of devices that are sensitive to contamination. In a zero-g environment, there is perhaps a greater tendency for contaminants to migrate from entrapped areas than would be the case in a one-g environment. Hence, special attention should be devoted to the elimination of features that may act as contaminant collection points and subsequent migration sources during zero-g operation.
It is believed that subsystem integration can best be achieved if a single vendor supplies the hardware to subsystem-level requirements. This approach contrasts with the lunar module stabilization and control subsystem procurement, which was at the assembly level with the lunar module contractor retaining integra- tion responsibility.
|Impedance mismatches in a radio-frequency (RF) electrical transmission line cause power loss and reflected energy.|
The impedance-mismatch problem between the gimbal drive actuator and the descent-engine control assembly resulted from a minor gimbal drive actuator change made without sufficient understanding of the interface effects. Although subsystem-level procurement is not a panacea for all problems, its appli- cation would make this sort of interface problem much easier to avoid than with assembly-level procurement.
REFERENCES/1/ Apollo News Reference, 1968
/4/ MECHANICAL DESIGN OF THE LUNAR MODULE DESCENT ENGINE
Jack M. Cherne, Manager,
Engineering Design Department,
Power Systems Division, TRW Systems,
Redondo Beach, California, U.S.A.
/5/ APOLLO EXPERIENCE REPORT DESCENT PROPULSION SYSTEM,
William R. Hammock, Jr.,
Eldon C. Currie, and
Arlie E. Fisher,
Manned Spacecraft Center, Houston, Texas 77058
/6/ APOLLO EXPERIENCE REPORT MISSION PLANNING FOR LUNAR MODULE
DESCENT AND ASCENT, Floyd V. Bennett, Manned Spacecraft Center
/7/ LMA790-2 - Lunar Module Vehicle Familiarization Manual - Nov 1, 1969
/8/ Don Eyles - Tales from the Lunar Module Guidance Computer, 2004
/9/ YouTube videos:
"LM - Capcom" audio and
"LM - Flight Control" audio.
/10/ NASA TN D-8227 -
M. D. Holley, W. L. Swingle, S. L. Bachman,
C. J. LeBlanc, H. T. Howard, and H. M. Biggs -
APOLLO EXPERIENCE REPORT - GUIDANCE AND CONTROL SYSTEMS:
NAVIGATION, AND CONTROL SYSTEM DEVELOPMENT PRIMARY GUIDANCE, May 1976
/11/ NASA TN D-8086 -
D. Harold Shelton -
APOLLO EXPERIENCE REPORT - GUIDANCE AND CONTROL SYSTEMS -
LUNAR MODULE STABILIZATION AND CONTROL SYSTEM, November 1975
/12/ NASA TN D-7990 -
Kurten, P. M.:
Apollo Experience Report - Guidance and Control Systems:
Lunar Module Abort Guidance System., 1975.
/14/ NASA TN D-7289
Willium H. Peters, Kenneth J. Cox
APOLLO EXPERIENCE REPORT - GUIDANCE AND CONTROL SYSTEMS -
DIGITAL AUTOPILOT DESIGN DEVELOPMENT, June 1973
/16/ Internet & Flicker
/17/ Lunar Module Structures Handout LM-5
/18/ George P. Sutton, Oscar Biblarz, "Rocket Propulsion Elements"
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