This section contains a discussion of the significant anomalies that occurred during the Apollo 15 mission. The discussion is divided into six major areas: command and service modules; lunar module; scientific instrument module experiments; Apollo lunar surface experiments package and associated equipment; government-furnished equipment; and the lunar roving vehicle.


14.1.1 Service Module Reaction Control System Propellant Isolation Valves Closed

During postinsertion checks, the quad B secondary isolation valve talkback indicated that the valve was closed, and the switch was cycled to open it. Subsequently, talkbacks indicated that both the primary and secondary valves for quad D were also closed, and these valves were reopened. At S-IVB separation (approximately 3 hours 22 minutes), all the aforementioned valves closed and were reopened. Upon jettisoning of the scientific instrument module door, the quad B secondary valve closed and was reopened.

This type of valve (a magnetic latching valve, shown in fig. 14-1) has, in previous missions, closed as a result of pyrotechnic shocks. Ground tests have shown that the valve will close at a shock level of approximately 80g sustained for 8 to 10 milliseconds. There were no indications of shock levels of the magnitude required to close the valve during launch.

Testing has shown that if a reversed voltage of 28 volts is applied to the valve, the latching voltage will drop to a point where the valve will no longer remain latched (magnet completely degaussed). In addition, at lower voltages with reversed polarity, the magnet would become partially degaussed.

During acceptance testing of one valve for command and service module 117, the latching voltage had changed from approximately 13 volts to 3 volts. Additional testing of the spacecraft 117 valve verified the low voltage condition. Additionally, the valve stroke was proper, thereby eliminating contamination as a possible cause of the problem. During the test, the valve was disconnected from spacecraft power (28 volts) and was being supplied power through a variable power supply (approximately 20 volts, maximum, applied to the valve). The valve was most likely subjected to a reversed polarity at a voltage level which would partially degauss the magnet. This may have been the cause of the valve closures during Apollo 15 launch phase.

A magnetic latching force test was not performed on the valves after assembly into the system for the Apollo 15 command and service module, as on some previous spacecraft. A test will be performed on subsequent assemblies to verify that the valve latching forces are acceptable.

This anomaly is closed.

14.1.2 Water Panel Chlorine Injection Port Leakage

Minor leakage was noted from the chlorination injection port when the cap was removed to perform the prelaunch water chlorination. The cap was reinstalled and the leak ceased. A leak of approximately 1 quart in 20 minutes also was noted at the chlorine injection port as the crew removed the injection port cap for the third injection at about 61 hours. The crew tightened the septum retention insert ( fig. 14-2) and satisfactorily stopped the leakage. Leakage recurred at about 204 1/2 hours and was corrected.

Postflight inspection and dimensional checks of the injection port assembly showed that all components were within established tolerances.

However, when the insert was tightened in accordance with the drawing requirements, the resulting septum compression was apparently insufficient to prevent the insert from loosening as a result of "O-ring drag" when the cap was removed. This allowed water leakage past the relaxed septums.

For future spacecraft, a shim will be installed under the insert shoulder to control the septum compression while allowing the installation torque to be increased to a range of 48 to 50 in-lb and, thus, preclude insert backout.

This anomaly is closed.

14.1.3 Service Propulsion System Thrust Light On Entry Monitor System

The service propulsion system thrust light located on the entry monitor system panel was illuminated shortly after transposition and docking with no engine firing command present. This light indicated the presence of a short to ground in the service propulsion system ignition circuitry. Ignition would have occurred if the engine had been armed.

The short was isolated to the system A delta-V thrust switch which was found to be intermittently shorted to ground ( fig. 14-3).

A test firing performed at 28:40:22 verified that the short existed on the ground side of the service propulsion system pilot valve solenoids.

The delta-V thrust switch ( fig. 14-4) was shorted to ground both before and after removal of panel 1 from the command module during postflight testing. After a change in panel position, the short-to-ground disappeared. The switch was then removed from the panel and X-rayed. The X- rays showed a wire strand extending from the braid strap which was thought to have caused the grounding problem. After switch dissection, an internal inspection verified that a strand extended from the braid strap; however, it did not appear to be long enough to cause a ground at any point within the switch(fig. 14-4). The bracket assemblies (pivot brackets, pigtail braids, and movable contacts) and the plastic liner were removed from the switch. Microscopic examination revealed that a wire strand (approximately 0.055 inch long) was present on the flange on terminal 2 (fig. 14-5). The strand appeared to be attached, but was later moved quite easily.

The bottom of the plastic case liner was examined, and showed no evidence of a scratch or deformation conforming to the shape of the wire strand. A sample wire strand was placed on a feed-through flange of a scrap switch unit, and the plastic case liner was pressed on as would occur during normal switch assembly. When the scrap switch was disassembled an indentation in the plastic case liner was readily apparent. This test indicated that the strand could not have been trapped between the case liner and the flange surface; therefore, it is postulated that it might have been enclosed in the cavity of feed-through terminal 2 ( fig. 14-5). The maximum clearance between the interior of the feed-through terminal wall and the terminal itself is 0.040 inch. A 0.055-inch-long wire strand could easily have bridged this distance, and yet is short enough to move quite freely within the feed-through terminal cavity. In fact, the strand subsequently fell into the cavity. Examination of the strand and cavity wall showed evidence of arcing. The strand could not be detected on the X-rays because that area was obscured by other poles in the switch.

Most of the switches on Apollo 16 and 17 spacecraft (3000 or 4000 series) required for crew safety or mission success were screened according to the following procedures.

The following switches for Apollo 16 were of an earlier series and have been replaced with 4000 series switches:

Switches required for crew safety and mission success for Apollo 17 which had not been screened according to the aforementioned procedures will also be replaced. In addition, two science utility power switches are to be disabled and stowed, and two circuit breakers are to be added to provide series protection for the command and service module/lunar module final separation function.

This anomaly is closed.

14.1.4 Integral Lighting Circuit Breaker Opened

The a-c bus 2 and the d-c bus B under-voltage alarms occurred and, subsequently, the integral lighting circuit breaker opened.

A short circuit sufficient to cause the circuit breaker to open would also cause the alarms. As a result of the problem, some display keyboard lights, the entry monitor system scroll lighting, and various other backlighting were not used for the remainder of the mission.
Postflight testing of the vehicle disclosed that the short circuit was in the mission timer. The timer was removed from the vehicle and returned to the vendor for further analysis. Teardown analysis revealed a shorted input filter capacitor.

The capacitor is rated for 200-volt d-c applications and is being used in an a-c application at voltages up to 115 volts. Since the dielectric in the ceramic capacitor is a piezolectric material (barium titanate), the 400-cycle a-c voltage actually causes the materials in the capacitor to mechanically vibrate at that frequency. Over a period of time, the unit could break down because of mechanical fatigue. This may have been the cause of failure of this capacitor.

There are two mission timers on the command module and one on the lunar module. The unit on the lunar module is separately fused. Fuses will be added to the units in the Apollo 16 and 17 command modules. Appropriate action will be taken to correct the timer design and an inline change will be made on both the command module and lunar module.

This anomaly is closed.

14-1.5 Battery Relay Bus Measurement Anomaly

At approximately 81-1/2 hours, the battery relay bus voltage telemetry measurement read 13.66 volts instead of the nominal 32 volts, as evidenced by battery bus voltage measurements. The crew verified that the same low voltage reading was present on the panel 101 systems test meter. When the crew moved the systems test meter switch, the reading returned to normal.

Postflight testing of the vehicle and all of the involved components revealed no anomalous condition which could have caused the problem but did isolate the problem to the instrumentation circuitry and verify that the functional operation of the bus was not impaired. Analysis indicates that the only way to duplicate the flight problem would be to connect a resistance of 2800 ohms from ground to the battery relay bus measurement circuit ( fig. 14-6). No resistance near this magnitude was measured during postflight testing. The most probable cause of the anomaly was that insulation resistance at the output terminal of the switch was lowered because of humidity.

This is the only time that a problem of this type has occurred during the Apollo Program and the probability of recurrence is considered to be very low. If the problem does occur again, other measurements are available for the determination of the battery relay bus voltage.

This anomaly is closed.

14.1.6 Mass Spectrometer Boom Talkback Indicated Half-Barberpole On Retract

The mass spectrometer boom did not fully retract on five of twelve occasions. Data analysis, supported by the crew debriefing, indicates that the boom probably retracted to within about 1 inch of full retraction. Cold soaking of the deployed boom and/or cable harness preceded each anomalous retraction. In each case, the boom retracted fully after warmup.

The deploy/retract talkback indicator is normally gray when off, when the boom is fully retracted, or when it is fully extended. The indicator is barberpole when the boom is extending or retracting, and will show half barberpole if the drive motor stalls. The crew noted this last condition on the incomplete retractions.

An inflight test of the Apollo 15 boom indicated that the problem was a function of temperature. Testing and examination of the Apollo 16 spacecraft showed that the failure was possibly caused by pinching of the cable harness during the last several inches of boom retraction. The cable could have been pinched between the bell housing and rear H-frame bearing ( figure 14-7), or a cable harness loop was jammed by a boom alignment finger against the bell housing ( fig. 14-8).

The mass spectrometer boom mechanism was qualified by similarity to the gamma ray boom mechanism. There are significant differences between the two designs and they are:

The differences between the two configurations are now considered to be significant enough to have required separate testing for the mass spectrometer boom assembly. Accordingly, a delta qualification test will be instituted and a thermal vacuum environmental acceptance test will be performed on each flight unit.

Additional failure modes revealed during the testing of the Apollo 16 unit are:

If the boom does not retract to within approximately 12 inches of full retraction, it will be jettisoned prior to the next service propulsion system firing. Tests have shown that the boom will not buckle during a service propulsion system firing when retracted to within 14.5 inches of full retraction.

Corrective actions for Apollo 16 are as follows:

This anomaly is closed.

14-1.7 Potable Water Tank Failure To Refill

The potable water tank quantity began to decrease during meal preparation at approximately 277 hours and failed to refill for the remainder of the flight. The waste water tank continued to fill normally and, apparently, accepted fuel cell water for this period. A similar occurrence had been noted earlier, at 13 1/2 hours, when the potable tank quantity decreased as the crew used the water, and remained constant until a waste water dump was performed at 28 1/2 hours. This decrease had been attributed to a closed potable tank inlet valve until the crew verified in their debriefing that the valve had been open during this time. The amount of water drained from the tank verified that the tank instrumentation was reading correctly.

During a postflight fill operation, with the waste tank inlet valve closed, and water introduced at the hydrogen separator, both the potable and waste water tanks filled.

The check valve between the fuel cell and waste tank dump leg (figure 14-9) was tested and found to leak excessively. A tear-down analysis of the check valve was performed and a piece of 300-series stainless steel wire (approximately 0.0085 by 0.14 inch) was found between the umbrella and the seating surface ( fig. 14-9). This contaminant could cause the umbrella to leak and yet move around sufficiently to allow adequate seating at other times. The wire most probably came from a welder's cleaning brush and was introduced into the system during buildup. Safety wires and tag wires are of a larger diameter than the one found. The check valve at the potable water tank inlet is of a different configuration and is spring loaded closed. The 1-psi pressure required to open this valve is a large pressure drop compared to the other components at the low flow of 1-1/2 lb/hour, and would, therefore, cause the water to flow to the waste tank.

The potable water tank inlet check valve was found to be contaminated with aluminum hydroxide, a corrosion product, of aluminum and the buffer. The potable water tank inlet nozzle was clean and free of corrosion. The check valve corrosion is not believed to have caused the problem, but could have contributed by increasing the crack pressure of the valve.

No corrective action is considered necessary since the contamination is considered to be an isolated case. If the problem should recur, the potable tank will start to fill when the waste tank is full.

This anomaly is closed.

14.1.8 Mission Timer Stopped

The panel 2 mission timer stopped at 124:47:37. Several attempts to start the clock by cycling the start /stop/reset switch from the stop to the start position failed ( fig. 14-10). The timer was reset to 124:59:00 using the hours, minutes, and seconds switches, and the timer again failed to start when the switch was cycled. The switch was then placed in the reset position. The timer reset to all zeros and started to count when the switch was placed in the start position. The timer was then set to the proper mission time using the hours, minutes, and seconds switches and operated properly for the remainder of the mission.

The timer and all associated equipment were still operating properly after the flight. Thermal, vacuum, and acceptance tests were performed and the cause of the failure could not be determined. Circuit analysis showed that the problem could be caused by one of five integrated circuits on the mounting board circuitry. These suspect components were removed and tested with negative results.

The failure was most probably caused by an intermittent problem within a component which later cured itself. If the problem occurs on a future mission and the timer will not restart, mission time can be obtained from the other timer in the command module, or from mission control. The failure would be a nuisance to the crew.

This anomaly is closed.

14.1.9 Main Parachute Collapse

One of the three main parachutes was deflated to approximately one fifth of its full diameter at about 6000 feet altitude. The command module descended in this configuration to landing. All three parachutes were disconnected and one good main parachute was recovered. Photographs of the descending spacecraft indicate that two or three of the six riser legs on the failed parachute were missing ( fig. 14-11).

Three areas that were considered as possible causes are:

Onboard and photographic data indicate that the forward heat shield - was about 720 feet below the spacecraft at the time of the failure. The failed link on the recovered parachute implies the possibility of a similar occurrence on the failed parachute. Based on parachute tow tests, however, more than one link would have had to fail to duplicate the flight problem. The two possible causes have been identified as hydrogen embrittlement or stress corrosion.

The command module reaction control system depletion firing was considered as a possible candidate because of the known susceptibility of the parachute material (nylon) to damage from the oxidizer. Also because the oxidizer depletion occurred about 3 seconds prior to the anomaly, and fuel was being expelled at the time the anomaly occurred ( fig. 14-13). Further, the orientation of the main parachutes over the command module placed the failed parachute in close proximity to the reaction control system roll engines. Testing of a command module reaction control system engine simulating the fuel (monomethyl hydrazine) dump mode through a hot engine resulted in the fuel burning profusely; therefore, the fuel dump is considered to be the most likely cause of the anomaly.

In order to eliminate critical processing operations from manufacture of the connector links, the material was changed from 4130 to Inconel 718.

Based on the low probability of contact and the minimum damage anticipated should contact occur, no corrective action will be implemented for the forward heat shield. Corrective actions for the reaction control system include landing with the propellants onboard for a normal landing, and biasing the propellant load to provide a slight excess of oxidizer. Thus, for low altitude abort land landing case, burning the propellants while on the parachutes will subject the parachutes to some acceptable oxidizer damage but, will eliminate the dangerous fuel burning condition. In addition, the time delay which inhibits the rapid propellant dump may be changed from 42 to 61 seconds. This could provide more assurance that the propellant will not have to be burned through the reaction control system engines in the event of a land landing. A detailed discussion of all analyses and tests is contained in a separate anomaly report (reference 7).

This anomaly is open.

14.1.10 Data Recorder Tape Deterioration

At about 240 hours, after over 100 tape dumps had been completed, the ground was unable to recover the data contained on about the first 20 feet of tape. To alleviate the problem, that portion of the tape was not used again.

An electrical and physical examination of the flight tape was conducted. Observation of the bi-phase output of the 51.2 kilobit pulse code modulated output from the playback showed a complete deterioration of the waveform for the first 20 seconds (12-1/2 feet), together with numerous complete dropouts. After 20 seconds, the bi-phase signal gradually improved to the point where, at 30 seconds, the signal appeared normal. The 64 kilobit pulse code modulated output was similarly affected to a lesser degree.

The first 30 feet of tape was scanned under magnifications ranging from 50X to 400X. Under 50X magnification, a distinct pattern of embedded particles could be observed ( fig. 14-14). The deposits were quite heavy over the first 12 feet of tape, and gradually tapered out so that, at 20 feet, very few particles could be observed. Under 400X magnification, individual flakes of deposited material were observed. The portion of figure 14-14 at 400X magnification shows a typical cluster of particles on the beginning portions of the tape.

A 10-foot leader coated with a silver oxide compound is spliced to the beginning and end of the magnetic tape roll to activate the end-of-reel sensors on the tape transport. There has been a history of this material flaking off and affecting tape performance. Tape screening procedures were implemented by the manufacturer in 1968 to eliminate this problem. No further problems were encountered until Apollo 15. The recording method for Apollo 14 and previous missions was considerably different than that for the Apollo 15 mission. Bit packing densities for the Apollo 15 mission tape approach 9000 bits per inch while those for the previous missions were only 800 bits per inch. Abnormalities in the tape would have considerably more effect with the higher packing density. The utilization of the Apollo 15 mission recorder is also considerably higher, allowing more time for deposits to build up.

An acceptance test (except for environmental verification) with a new tape was conducted on the flight recorder and all parameters were within specification with little change in absolute values from the pre-delivery test.

Inspection on the magnetic heads under 20X magnification disclosed four scratches, one of which is shown in figure 14-15. An overlay was made of the scratches with respect to the accumulation of silver oxide on the tape; two of the four scratches aligned perfectly with the silver oxide accumulation. The scratches must have scraped loose the silver oxide on the leader. Operation of the recorder would then distribute the silver oxide particles along the tape. During the manufacture of the Apollo 16 recorder, it was discovered that the heads were being scratched by handling. The Apollo 15 recorder heads were probably also scratched during manufacture. The scratches would not have been detected during acceptance inspection since they are not visible at the 7X magnification used during that inspection.

Removable head covers have been provided to protect the heads from handling damage when the recorder covers are not installed. These covers have been used since early in the buildup cycle of the Apollo 16 and 17 data recorders. The recorder heads have been examined under 20X magnification and no scratches were found.

This anomaly is closed.

14.1.3-1 Digital Event Timer Obscured

The seconds digit of the digital event timer, located on panel 1, became obscured by a powder-like substance that formed on the inside of the glass. Postflight analysis of the unit disclosed that the substance on the window was paint which had been scraped from the number wheel by the idler gear. The idler gear is free to rotate on the shaft ( fig. 14-16); however, the design allows the stainless-steel shaft to also rotate. The stainless steel shaft bearing points are in the magnesium motor plate and the shaft rotation wears away the softer magnesium material.

Inspection of the unit showed that the magnesium bearing points had been elongated as shown in figure 14-17. Torque from the stepping motor applied to the idler gear not only resulted in rotation of the shaft but also caused the shaft to tilt ( fig. 14-17). The wearing eventually allowed the shaft to tilt sufficiently to cause the gear to rub against the number wheel. 'When the timer counted down, the motor torque threw the gear teeth into the front edge of the counter wheel. Testing indicates that this bearing hole elongation occurs after approximately 500 hours running time (specification life is 1400 hours).

A review of the history of the unit shows that it was built in 1966. Prior to installation in Apollo 15, the unit was modified because of failures on other timers. Brass shims (fig. 14-17) were installed to prevent the idler gear from rubbing on the number wheel.

The analysis of those failures revealed that the idler gear was rubbing paint off the number wheel and paint particles prevented the slip rings and brushes from making good contact. A review of drawing tolerances showed that an interference could occur and the addition of the shims appeared to be adequate corrective action. These failure analyses did not reveal the problem of the elongation of the bearing points since it is not obvious until the timers are disassembled.

Units for future flights will be visually inspected by looking through the window for paint flakes and signs of wear.

This anomaly is closed.

14.1.12 Crew Restraint Harness Came Apart

The restraint harness on the right side of both the center and right crew couches came apart during lunar orbit. The assemblies had become unscrewed, but the crew was able to retrieve all the parts except one cap and reassemble the harnesses satisfactorily for landing. The mating plug for the missing cap was held in place with tape.

The plug-and-cap assembly ( fig. 14-18), which is part of the universal assembly that attaches the restraint harness to the couch seatpan, separated. (There are a total of six plug-and-cap assemblies on the crew couch, two per man.) The plug component (bolt) has a nylon insert in the threaded portion that acts as a locking device. Back-and-forth rotation of the adjuster link can cause the plug-and-cap assembly to unscrew from each other. Checks on the four other Apollo 15 assemblies showed zero torque on two of the units and minimum specification value (2.0 in-lb) on the others. The loss of torque is apparently due to cold flow of the plastic self-locking pellet, causing a loss of friction against the mating threads.

A thread locking sealant will be used to prevent the problem on future missions.

This anomaly is closed.

14.1-13 Loose Object In Cabin Fans

During portions of the flight when the cabin fans (fig. 14-19) were activated, the crew heard sounds like an object striking the blades. Cycling the fans several times allowed the object to be retained in a position that precluded it from interfering with fan operation.

Inspection of the fans revealed considerable gouging on the leading edges of the blades of both fans ( fig. 14-19). No marks were found on the outlet de-swirl vanes of either fan. After extensive examination, including the lunar dust filter, a 1/4-inch washer was found in the ducting.

Although the fan inlet is protected by screens in addition to the heat exchanger core, the outlet is relatively open and the washer could have drifted in and out when the fans were not operating. The outlet was protected only during the limited time when the lunar dust filter was installed. The washer could also have been left in the duct during assembly.

No hardware changes are contemplated. Should the anomaly occur on a subsequent flight, no detrimental effects would result.

This anomaly is closed.

14.1.14 Scanning Telescope Visibility

The crew reported that excessive attenuation of light through the scanning telescope existed throughout the flight. The telescope was adequate to perform landmark tracking while in lunar orbit, but the crew was unable to identify constellations, even though large numbers of stars could be seen by looking out the spacecraft window.

Visual observations through the telescope ( fig. 14-20) were made at the spacecraft manufacturer's facility, and no degradation could be observed. A luminescent transmittance test was performed on the telescope before removal from the spacecraft and transmittance was calculated to be 25 percent. This compares with an acceptance test value of 55 percent. The decrease is due to the entry environment and sea water contamination. The 30-percent decrease agrees well with the expected results and is not significant as far as being able to see and recognize constellations is concerned. For comparison, the earth's atmosphere normally causes a 50 percent loss in star intensity; therefore, observing stars from earth with a telescope with a 50-percent transmittance would be equivalent to observing stars from a spacecraft in flight using a telescope with a 25- percent transmittance.

The flight anomaly was reproduced in the laboratory by placing the optical unit assembly, the removable eye piece, and the optics panel in a chamber wherein the environmental conditions that existed in the cabin during flight were duplicated. Condensation on the eyepiece window and, to a lesser extent, on the prisms in the removable eyepiece caused the transmittance to decrease to about 4 percent.

A heater will be added to the removable eyepiece to prevent fogging in the eyepiece assembly and on the eyepiece window.

This anomaly is closed.

14.1-15 Gyro Display Coupler Roll Alignment

The crew reported that the roll axis did not align properly when the gyro display alignment pushbutton was pressed. The roll axis error was not nulled, whereas, the pitch and yaw axes were. Only by depressing the align pushbutton for progressively longer periods, and eventually, by moving the roll-axis thumbwheel, could the roll error be nulled.

For normal operation during alignment, resolvers in the gyro display coupler electronics are compared to resolvers, one for each axis, on the thumbwheels used to set desired attitude. The difference is an error signal. The error is displayed on the attitude error needles, and the signal is used to drive the resolvers to match the attitude set on the thumbwheels. The anomaly could have been caused by either of two failure modes. An intermittent open in the roll axis align loop or a low gain problem in the electronics ( fig. 14-21).

The gyro display coupler and attitude set control panel were put into the hardware evaluator simulator and functionally tested at the systems level in the actual spacecraft configuration in an attempt to repeat the flight problem. During this testing, an out-of-tolerance condition was observed on the attitude set control panel. This condition could have caused a gain type problem and been the cause of the flight anomaly. The measured resistance of the thumb wheel resolvers increased from the nominal in all three axes by as much as 1000 ohms. Normally, this value does not vary by more than 1 ohm. In order for this condition to have been the cause of the anomaly, a resistance change in the roll axis would have had to be an order of magnitude larger than that measured postflight. The resistance change is caused by contamination between the slip rings and the thumbwheel resolvers. As a result of the flight anomaly, several resolvers were examined and contamination was detected. The corrective action is to wipe the resolvers clean by rotating them several hundred revolutions. The attitude set control panels in Apollo 16 and 17 will be checked and the resolvers will be wiped clean, or will be replaced if necessary.

The flight condition could also have been caused by either of two golden-g" relays failing to close. Two failure modes have been determined. One failure mode is "normally open-failure to close", and other, "normally closed - failure to open", both caused by contamination. "Golden-g" relays were the subject of an extensive review in 1966 and 1967. Relays were classified as (1) critical, (2) of major importance, (3) of minor importance, and (4) having no effect. It was decided at that time to (1) make critical relays redundant, (2) improve screening tests, and (3) take no corrective action for non-critical relays. Both of the suspect relays are of major importance in that either one would cause loss of the normal alignment capability of the backup attitude reference system. The attitude reference system could be aligned, but extensive work-around procedures would have to be used.

Tests performed on the roll axis align enabling relay revealed contamination which could have caused the flight anomaly. The rationale developed during the "golden-g" relay review is applicable at this time.

This anomaly is closed.

14.1.16 Unable To Open Circuit Breaker Supplying Main A Power To Battery Charger

The circuit breaker tying the battery charger to main bus A could not be opened manually during postflight testing. This breaker was not required to be opened during the flight.

A green residue on the aluminum indicator stem at the copper mounting bushing jammed the stem and prevented operation. Some of the residue was removed for chemical analysis. The rest of the residue was dissolved by the application of distilled water, thereby freeing the breaker. The green residue was predominantly sodium-copper carbonate hydrate. Traces of sodium chloride, and other metals were also present. These products would result from salt water corrosion. Salt water could have been introduced by sea water splashing on the breaker after landing or by urine or perspiration released in the cabin during flight.

No corrective action is considered necessary.

This anomaly is closed.

14.1-17 Pivot Pin Failure On Main Oxygen Regulator Shutoff Valve

The toggle-arm pivot pin for the side-A shutoff valve of the main oxygen regulator was found sheared during postflight testing. With the pin failed, the shutoff valve is inoperative in the closed position, thus preventing oxygen flow to the regulator.

The pivot pin attaches the toggle arm to the cam holder and is retained in place by the valve housing when properly assembled ( fig. 14-22).

Failure analysis showed that the pivot pin failed in single shear and bending. This failure resulted from improper shimming which allowed the pivot pin to come out of one side of the cam holder as shown in figure 14- 22. Analysis and testing has shown that the pin strength is adequate in double shear, but will fail in single shear and bending with a force of about 70 pounds applied at the tip of the toggle arm when it is in the closed position. No marks were found on the toggle arm to indicate that it had been struck by some object.

Inspection criteria to assure that valves now installed in other spacecraft are properly assembled have been developed from a study of adverse tolerance buildups associated with the valve components. These criteria are that the lock nut does not protrude and the number of shims does not exceed six (fig. 14-22).

This anomaly is closed.

14.1.18 Crew Optical Alignment Sight Fell Off Stowage Mount

The crew optical alignment sight fell from its stowage mount during landing because the locking pin which secures it was not engaged. Normally, when the sight is placed into the mount, the locking pin is raised automatically by a ramp and the pin is moved into the locking pin hole by spring action ( fig. 14-23). Postflight examination showed that the ramp had been gouged preventing raising of the pin by the ramp. The cause of the gouge is not known.

The crew optical alignment sights for Apollo 16 and future spacecraft will be fit-checked to insure proper operation of the latching mechanism. Also, the Apollo Operations Handbook and crew checklist are being revised to include verification of the latching pin engagement prior to entry.

This anomaly is closed.


14.2.1 Water/Glycol Pump Differential Pressure Fluctuations

Variations were noted in the water/glycol pump differential pressure shortly after the cabin depressurizations for the standup extravehicular activity and the second extravehicular activity. The variations were similar on both occasions. The pressure differential decreased from about 20 psid (normal) to about 15 psid, then increased to about 27 psid and returned to normal ( fig. 14-24).

The total times for the cycles were 3 minutes during the standup extravehicular activity and 10 minutes during the second extravehicular activity. The pump discharge pressure remained relatively stable throughout both periods. If pressure fluctuations had taken place in the heat transport system, both the pump discharge pressure and differential pressure should have varied together.

After the second fluctuation occurred, water/glycol pump 2 was selected because of the erratic differential pressure. All parameters were normal during pump 2 operation. The crew reselected pump 1 prior to egress for the second extravehicular activity, and it operated satisfactorily. Later, after docking, the pump was turned off momentarily and the pump discharge pressure readout was verified as correct since it decreased to the accumulator pressure of 7.8 psia.

The lunar module cabin humidity was high at the initial manning for descent because the command module cabin humidity was high. Furthermore, a water spill in the lunar module cabin after the first extravehicular activity again produced high humidity. Consequently, water would have condensed on the cold 1/8-inch water/glycol sense lines between the pump assembly and the pressure transducer ( figs. 14-25 and 14-26), and the water would have frozen and sublimed at the next cabin depressurization. Since there is no flow in the sense lines, as little as 0.002 inch of condensed moisture on the outside of these lines would freeze the fluid and cause the fluctuations in the indicated differential pressure, but would not affect system operation. Consequently, no corrective action is required.

This anomaly is closed.

14.2.2 Water Separator Speed Decrease

The speed of water separator 1 decreased to below 800 rpm and tripped the master alarm during the cabin depressurization for the standup extravehicular activity. Separator 2 was selected and operated properly at approximately 2400 rpm. After approximately 1 hour of separator 2 operation, separator 1 was reselected and performed satisfactorily throughout the remainder of the mission.

Cabin atmosphere is cooled and passed through one of the water separators (fig. 14-27) where condensed water is separated by centrifugal force and picked up by a pitot tube. The water is then piped from the pitot tube, through a check valve, to the water management system where it is used in the sublimator.

Cabin humidity was high before the standup extravehicular activity and, because the water and structure were cold, the line between the water separator pitot tube and the water management system was cold. Under these conditions, water would condense on the outside of the line ( fig. 14-27), and when the cabin was depressurized for the standup extravehicular activity, the water on the outside of the line would boil, freeze, and sublime, thereby freezing the water in the line.

Analysis shows that as little as a 0.01-inch film of water on the outside of the line will freeze the line at cabin depressurization. Freezing within the line will cause the separator to slow down and stall because of excessive water. Separator 2 had not been used. Therefore, no water was in its outlet line to freeze. Consequently, it operated successfully when activated.

Analysis and tests have shown that freezing of the line will not damage any spacecraft hardware. If such freezing occurs, the other separator would be used. Therefore, no corrective action is required.

This anomaly is closed.

14.2.3 Broken Water Gun/Bacteria Filter Quick Disconnect

A water leak occurred at the quick disconnect between the bacteria filter and the water gun (fig. 14-28) shortly after the first extravehicular activity. The leak was caused by the plastic portion of the disconnect being broken and was stopped by removing the filter.

When the water gun/bacteria filter combination is properly stowed, the gun is held in a U-shaped boot ( fig. 14-28). Both the gun and filter are held in place by straps with the hose protruding above the liquid cooling assembly. Bending the hose can easily exceed the force required to break the quick disconnect between the bacteria filter and the water gun if the filter is not strapped down. A test showed that a torque of 204.6 inch-pounds caused a similar quick disconnect to break in half.

Based on the torque value of 204.6 inch-pounds, a force of approximately 27 pounds applied at the hose quick disconnect interface would break the disconnect between the filter and the water gun.

The plastic parts will be replaced by steel inserts in all locations where the quick disconnects are used in the lunar module and command module.

This anomaly is closed.

14.2.4 Intermittent Steerable Antenna Operation

Random oscillations occurred while the steerable antenna was in the auto-track mode. The oscillations were small, and damped without losing auto-track capability ( fig. 14-29). The three times when the oscillations became divergent are shown in figure 14-30.

All oscillations indicated characteristics similar to the conditions experienced on Apollo 14 ( fig. 14-31). After Apollo 14, one of the prime candidates considered as a possible cause was incidental amplitude modulation on the uplink signal. A monitor capable of detecting very small values of incidental amplitude modulation was installed at the Manned Space Flight Network Madrid site for Apollo 15. The data from this monitor indicated that no amplitude modulation existed on the uplink at the frequencies critical to antenna stability during the problem times. Consequently, incidental amplitude modulation has been eliminated as a possible cause of the antenna oscillations, and the problem must be in the spacecraft.

There are two possible causes of the problem in the spacecraft. The first is that electrical interference generated in some other spacecraft equipment is coupled into the antenna tracking loop. A test performed on the Apollo 16 lunar module showed that no significant noise is coupled into the tracking loop during operation of any other spacecraft equipment in the ground environment. The second is that temperature or temperature gradients cause an intermittent condition in the antenna which results in the oscillations. The acceptance tests of the antennas do not include operation over the entire range of environmental temperatures; therefore, a special test is being conducted on a flight spare antenna to determine its capability to operate over the entire range of temperatures and gradients.

This anomaly is open.

14.2.5 Descent Engine Control Assembly Circuit Breaker Open

The descent engine control assembly circuit breaker was found open during the engine throttle check after lunar module separation from the command and service module. The circuit breaker was closed and the check was successfully performed.

If there was a short circuit condition, it is highly unlikely that the fault would have cleared at the instant the breaker opened. Thus, when the breaker was reset, it would have reopened. Data were reviewed for current surges large enough to trip the 20-ampere breaker, but none were found. The crew stated that they may have left the circuit breaker open. This is the most probable cause of the anomaly.

This anomaly is closed.

14.2.6 Abort Guidance System Warning

Abort guidance system warnings and master alarms occurred right after insertion into lunar orbit and at acquisition of signal prior to lunar module deorbit. The first one was reset by the crew; the second persisted until lunar impact. Performance of the abort guidance system appeared normal before, during, and after the time of the alarms.

Exceeding any of the following three conditions in the abort guidance system can cause the system warning light to illuminate. In each case, the warning light would reset automatically when the out-of-tolerance condition disappears.

A fourth condition which could cause the warning light to illuminate is the receipt of a test-mode fail signal from the computer. This condition requires manually resetting the warning light by placing the oxygen/water quantity monitor switch to the CW/RESET position.

The conditions for generating a test mode fail signal are as follows (none of these conditions was indicated in any of the computer data).

Computer routines running at the time of the warning have a worst-case execution time of 18.425 milliseconds of the allowable 20 milliseconds; therefore, a timing problem should not have occurred.

After the warning at insertion, the crew read out the contents of the computer self test address 412, but there was no indication of a test-mode fail. The crew did not, however, reload all zeros into address 412 as is required to reset the flip-flop (fig. 14-32) which controls the test-mode fail output in the computer. Consequently, a second test-mode fail would not have caused an abort guidance system warning. The fact that a second warning did occur restricts the location of the failure to the output circuit of the computer, the signal conditioner electronics assembly, or the caution and warning system.

The abort electronics assembly test mode fail output drives a buffer ( fig. 14-32) in the signal conditioner electronics assembly which, in turn supplies the test mode fail signal to the caution and warning system. The test mode fail buffer is the only application in the lunar module where a transistor supplies the input signal to the buffer and where the low side of the input to the buffer is not grounded. Tests have shown that some buffers feed back 10-kHz and 600-kHz signals to the buffer input lines. These signals are completely suppressed when the low side of the buffer input is grounded.

With the noise signal voltages present, the test mode fail driver is more susceptible to electromagnetic interference. An analysis indicates that,, with worst-case noise signals present, an induced voltage spike of only about 3.5 volts will cause the test mode fail driver to momentarily turn on and latch the caution and warning system master alarm and abort guidance system warning light on.

For future spacecraft, the low side of the input to the buffer will be grounded.

This anomaly is closed.

14.2.7 No Crosspointer Indication

There was no line-of-sight rate data on the Commander's crosspointers during the braking phase of rendezvous. The existence of line-of-sight rates was verified by observing the position of the command module relative to the lunar module. The scale switch was in the low position; however, none of the other switch positions was verified. The power fail light was off, indicating that the Commander's crosspointer circuit breaker was closed.

The rate error monitor switch is used to select either rendezvous radar data or data selected by the mode select switch for display. The mode select switch selects velocities from the primary guidance system, the abort guidance system, or the landing radar for display. No telemetry data are available on the position of the rate error monitor switch and the mode select switch, and the Commander reported that he did not look at the flight director attitude indicator.

Four possible conditions could have affected the display of radar antenna rate data ( fig. 14-33).

The most probable failure is an open in the signal return line. Information could still be deduced from antenna position data which is displayed on the flight director attitude indicators. Ground tests and checkout may not show this type of failure. If the open is temperature sensitive, a complete vehicle test involving vacuum, temperature, and temperature gradients would be required to insure that failures of this type would not occur in the flight environment. This type of testing is not practical at the vehicle level. Consequently, no corrective action is planned.

This anomaly is closed.

14.2.8 Broken Range/Range Rate Meter Window

Sometime prior to ingress into the lunar module, the window of the range/range-rate meter broke ( fig. 14-34). Upon ingress, the crew saw many glass particles floating in the spacecraft, presenting a hazardous situation.

The window is an integral part of the meter case and is made of annealed soda-lime glass 0.085-inch thick. The meter case is hermetically sealed and pressurized with helium to 14.7 psia at ambient temperature. At the maximum meter operating temperature and with the cabin at vacuum, the pressure differential can be as high as 16.1 psi. This pressure differential is equivalent to a stress level of 6680 psi in the glass.

Glass will break when there is a surface flaw large enough to grow at the stress levels present. The threshold flaw size in a dry environment is about the same as the critical flaw size and immediate breakage occurs. The critical flaw size remains the same in a humid environment, but the threshold flaw is much smaller. For annealed soda-lime glass at a stress level of 6680 psi, the critical flaw depth is 0.0036 inch, and for a humid environment, the threshold flaw depth is 0.000105 inch.

A surface flaw deeper than the threshold depth for the glass operating stress must have existed on the outside of the meter window at launch. The flaw started growing as the cabin depressurized during the launch phase, and finally grew large enough for the glass to break.

For future missions, an exterior glass doubler will be added to the flaw depth to 0.0036 inch and the critical flaw depth to 0.032 inch. This should prevent future fatigue failures since there are no reported fatigue failures in soda-lime glass at stresses below 2000 psi. In addition, all similar glass applications in the lunar module and command module were reviewed and changes are being made. In the command module, transparent Teflon shields will be installed on the:

In the first three above applications, the shields will be held in place with Velcro and will be installed only when the cabin is to be depressurized. The shield on the entry roll monitor indicator will be permanently installed.

In the lunar module, tape will be added to the flight director attitude indicators and an exterior glass shield will be installed over the crosspointers to retain glass particles. The data entry and display assembly window was previously taped to retain glass particles.

This anomaly is closed.


14.3.1 Panoramic Camera Velocity/Altitude Sensor Erratic

Telemetry received from the first panoramic camera pass on revolution 4 indicated that the velocity/altitude sensor ( fig. 14-35) was not operating correctly.

The velocity/altitude sensor measures the angular rate of travel of the spacecraft relative to the lunar surface. The sensor output is used to control the cycling rate of the camera, the forward motion compensation, and the exposure. The sensor normally operates in the range of 45 to 80 miles altitude. If, at any time, the indicated velocity/altitude is out of this range, the sensor automatically resets to the nominal value of 60 miles. The sensor operated properly for brief periods of time, but would drift off-scale high (saturate), and then reset to the nominal value corresponding to a 60-mile altitude.

Breadboard tests and circuit analyses of the velocity/altitude electronics ( fig. 14-36) did not indicate failure. Tests were conducted in which endless belts of lunar scene photography from Apollo 8 and 15 were passed in front of velocity/altitude sensors. Sensors from the prototype and qualification units, and flight unit number 1 were used. By varying the illumination level, sensor performance somewhat similar to the Apollo 15 anomaly could be obtained.

The results of the tests, coupled with analyses of the basic sensor design, indicate that the problem is related to the optical signal-to-noise ratio. The remaining flight hardware will be modified to improve this ratio. The optical signal will be enhanced by increasing the lens aperture from f/4.0 to f/3-5 and by deleting the infrared filter. The optical noise (reflections) will be reduced by increasing the length of the lens hood and by repositioning the sensor so that the camera's forward plume shield will not be in the field of view of the sensor. In addition, a manual override of the velocity/altitude sensor will be provided on the remaining flight units. By using a three-position switch, two preselected velocity/altitude ratios will be provided, as well as the automatic function.

This anomaly is closed.

14.3.2 Loss of Laser Altimeter Altitude Data

The laser altimeter exhibited two anomalous conditions during the mission:

The photomultiplier tube power supply anomaly was duplicated when a relay which had been removed from a flight unit because it had an audible "buzz" was installed in the prototype altimeter. The relay serves no function in flight, but is a safety precaution for ground personnel working on the altimeter ( fig. 14-38). The relay contacts close when the altimeter is turned off, discharging the high voltage stored in the pulse forming network capacitors.

It is suspected that the audible "buzz" is accompanied by electromagnetic interference that is coupled into the video amplifier in the laser receiver ( fig. 14-39). The video amplifier is a principal element in the automatic gain control circuit which controls the output of the photomultiplier tube power supply. The electromagnetic interference from the relay can thereby result in the automatic gain control holding the power supply in the idling mode until the pulse forming network is discharged in firing the laser. The relay and resistors that comprise the bleed-down circuit will be removed from the remaining flight altimeters.

The cause of the low output power anomaly has not been isolated. A review of the manufacturing records has established that the flight unit was the same as the qualification unit with regard to parts, processes, and manufacturing methods. Investigations indicate that the fault most likely occurred in the laser module.

An automatic power compensation circuit will be incorporated into the remaining flight units. The circuit will increase the pulse forming network voltage by about 50 volts each time the laser power falls below an established threshold value as sensed by a photodiode. Design feasibility tests have been completed on a breadboard circuit. The results show that the circuit will maintain the power output at a level sufficient to provide proper ranging.

This anomaly is closed.

14-3.3 Slow Deployment Of Mapping Camera

The extension and retraction times of the deployment mechanism subsequent to the first extend/retract cycle were two to three times longer than the preflight nominal time of approximately 1 minute 20 seconds. Also, the camera could not be fully retracted after the final deployment. During the transearth extravehicular activity, an inspection of the mapping camera and associated equipment showed no evidence of dragging or interference between the camera and the spacecraft structure, the camera covers, or the cabling.

The first extend and retract cycle times were 1 minute 20 seconds and 1 minute 17 seconds, respectively. The second retraction required 2 minutes 30 seconds and the third retraction and fourth extension required slightly more than 4 minutes. The second and third extensions occurred while the telemetry, system was in the low-bit-rate mode; therefore, these deployment times are not obtainable. Subsequent extensions and retractions required 2 to 4 minutes.

Load tests show that a restraining force of 250 pounds would increase the deployment time to 1 minute 45 seconds. With one of the two extend/retract mechanism motors operating, the 250-pound restraint would increase the deployment time to 2 minutes 25 seconds.

Voltage tests show that 12 volts to the motors (28 volts dc nominal rating) would result in deployment times of approximately 4 minutes. Had this occurred during the mission, however, the indicator which shows that power is applied to the motors would have displayed a partial barberpole during deployment operations. The barberpole indicator is connected in parallel with the motors and, since the position is voltage-dependent, it can be used to approximate the voltage levels to the motors. During the flight, a full barberpole indication was always observed.

Apparently, the problem first occurred sometime between the first and second retractions. During this period, a 4-second service propulsion system firing was performed for lunar orbit circularization. An evaluation of vibration test data indicates, however, that the circularization firing was probably not a factor in the anomaly. An investigation is being made to determine if there is mechanical interference between the camera and the reaction control system plume protection covers.

This anomaly is open.

14-3.4 Gamma Ray Spectrometer Calibration Shifts

During the mission, the gamma ray spectrometer experienced a downward gain shift of approximately 30 percent, but this was compensated for by commanding the high-voltage step function from the command module. The drift decreased with time at an initial rate of 1 percent per hour and a final rate of 0.4 percent per day. Near the end of the mission, the gamma ray spectrometer was operating in a relatively stable state at 824.8 volts (high voltage step 6). (A step 4 voltage of 777.8 volts was the normal position in preflight operation.) The spectrometer to be flown on Apollo 16 was aged at flux rates representative of those encountered in lunar operation. The unit has stabilized after having experienced a gain change of approximately 8 percent.

After transearth injection, a temporary eight-channel zero reference shift was observed. This shift disappeared when the instrument was repowered after the transearth extravehicular activity, and subsequent instrument operation was normal for about 25 hours during transearth coast. Shortly before entry, the offset shift reappeared and remained until the experiment was turned off. Normalization of the data during processing will compensate for this offset.

Tests conducted with the qualification unit verified that the change in gain was due to aging effects of the photomultiplier tube in the gamma ray detector assembly as a result of high cosmic ray flux rates in lunar operation. The zero shift appears to be associated with the run-down inhibit signal between the clock-gate module and the analog-to-digital converter ( fig. 14-40).

Absence of this signal at a particular point in the analog-to-digital converter removes a 3-microsecond offset in the pulse height analyzer. The resulting effect is an overall eight-channel offset in the spectrum. The qualification unit was partially disassembled and tests showed that either an open or a shorted wire within the pulse height, analyzer can result in an eight-channel offset. An inspection of the circuit in the qualification unit disclosed no design deficiency which would cause this type of failure. Since the eight channel zero offset does not significantly impact overall data quality, no corrective action is contemplated.

This anomaly is closed.


14.4.1 Problems During The Lunar Surface Drilling Operations

The Apollo lunar surface drill performed well during the lunar surface activities; however, the following problems related to drilling operations were encountered: Difficulty in penetrating the surface to the desired depth with the bore stems.- Although the average penetration rate for the two bore stem holes was reasonable (approximately 120 inches per minute for hole 1, and 18 inches per minute for hole 2), it was necessary to stop both holes at approximately 60 percent of the depth desired.

The bore stem sections are made of a fiberglass and boron filament laminate, chosen for its optimum thermal characteristics as a casing for the heat flow experiment probe. The sections are approximately 21 inches long with tapered male and female joints. One-inch double-thread spiral lead flutes are provided on the exterior surface to transport the soil chips from the drilled hole to the surface. The depth of the flutes is 0.050 inch for about 18 inches, but the flutes almost disappear at the joint area where the wall thickness must follow the taper of the joint ( fig. 14-41). As a result, the volume of chip flow to the surface is slowed considerably in looser soil formations, and stalled by the packing of the chips in high-density formations.

In order to reduce the time required, prevent damage to bore stems, and increase the probability of attaining the full depth, the following modifications and corrective actions are being implemented:

This anomaly is closed. Difficulty in releasing bore stems from drill adapter. Use of the normal procedure for releasing the lunar surface drill head from the bore stems was hampered by the bore stems turning freely in the lunar soil.

In the bore stem drilling position, the key blocks are restrained inside, and the spindle drives against the shoulder of the adapter outer shell ( fig. 14-42). The operational sequence to release the bore stem from the adapter includes the following steps:

In development ground tests, the soil friction usually kept the bore stem from turning in this operation. When there was insufficient friction from the soil, the bore stem was grasped with the gloved hand. On Apollo 15, the soil did not hold the bore stem, and the core-stem wrench was used to hold the bore stems for this operation.

The single-purpose core stem wrench is fitted to the 0.983-inch- diameter titanium core stem, but the 1.088-inch throat will admit the 1.075inch-diameter bore stem. The softer bore stem (boron-fiberglass laminate) can be deformed and present some difficulty in wrench removal, with possible damage to the bore stem. A wrench to fit both bore and core stems will be provided.

The change of the bore stem joint, discussed in the previous anomaly, will result in the elimination of the present bore stem drill adapter. The bore stem will thread into the power head spindle adapter in the same manner as the core stems on Apollo 15 and a spindle thread reducer will be provided to fit the core stems. In addition, the training models and procedures will by updated to reflect equipment changes.

This anomaly is closed. Bore stem damage near the first joint. The probe would not go to the bottom plug of the bottom bore stem in hole 2, but stopped at a point about 6 inches above the first joint. Examination of photographs and heat flow probe data indicate that, near the end of the bore stem drilling operation, the first joint was separated when the drill and drill string were moved vertically (up and down) in an attempt to improve the drill penetration rate. Easier penetration (for approximately 6 inches) was reported by the crew, but it resulted from the bottom of the second section apparently performing more in a coring manner ( fig. 14-43) with the lunar soil entering the second section of the bore stem.

The change from boron/fiberglass to threaded titanium in the bore stem joint will prevent a repetition of such a separation.

This anomaly is closed. Difficulty in core stem removal from the drilled hole in the lunar surface.- Friction of the compacted soil in the drill flutes can build up substantial forces against core stem removal in a deep hole in some soil formations. This was illustrated in premission and drill development experiences.

Interference from the compacted material in the drill flutes can be reduced and core stem removal eased by pulsing the power head when at the bottom of the hole without upward and downward motion of the drill stem. Ground tests have indicated that the best results are obtained when the power head is pulsed just before the power head is removed to add each core stem section. The tendency to auger, as reported by the crew, is also reduced by pulsing the power head before each new core stem is added.

To assure maximum core return and minimum core disturbance for this mission, and without having the benefit of some of the experience from later ground tests, the crew did not pulse the power head. In addition, the core stem string was left in the ground for several hours before the crew returned for its final removal. The core stem string was removed with considerable physical effort, but a very complete core was recovered.

A mechanical assist (modified jacking mechanism) will be mounted on the treadle for easier core removal from difficult formations. Training and procedural changes will be implemented so that the drill motor will be pulsed before the addition of each core stem.

This anomaly is closed.

14.4-1.5 Difficulty in separation of core stem sections.- The sections of the core stem string could not be separated using the vise and wrench because the vise had been mounted on the pallet backward. The six section core stem string was removed from the core hole as a single unit and brought to the vise on the lunar roving vehicle. Three sections were separated individually with hand friction on one side of the joint and the wrench on the other side. The remaining three sections were returned to the earth in one piece.

The configuration of the core stem vise is the same as that of the core stem wrench head. The vise is mounted on a bracket on the lunar roving vehicle aft chassis pallet, located on the right hand side of the vehicle. The core stem wrench head is similar to the conventional pipe wrench head, with one fixed jaw and one pivoted jaw. The throat width is not adjustable and is designed to fit the outside diameter of the core stem.

As mounted, the vise would hold the core stem so that the joint could be tightened by rotating the wrench on the adjoining section. However, the vise would not hold in the opposite direction so that the joint could be loosened and separated ( fig. 14-44). Working on the inboard side of the vise, the core stem could have been held properly for loosening; however, there is insufficient clearance on the inboard side of the vise for wrench rotation, and the distance to the other side of the lunar roving vehicle is greater than the length of a core stem section.

The installation drawing of the vise was in error and has been corrected to assure correct orientation of the vise for Apollo 16. The training vise was installed backward from the erroneous drawing, but correct for loosening the stems.

This anomaly is closed.

14.4.2 Central Station Rear Curtain Retainer Removal Lanyard Broke

To remove the retainer for the central station rear curtain, added for Apollo 15, it was necessary to remove two retaining pins ( fig.14-45). The two pins, a universal handling tool fitting, and the curtain retainer are joined by a three-section lanyard. When the universal handling tool was inserted in the fitting and raised to remove the first pin, that section of the lanyard broke. When an effort was made to remove both pins simultaneously by inserting the handle under the lanyard joining the two pins, that part of the lanyard broke. The pins and retainer were then removed by hand.

The Dacron lanyard is being changed from a 50-pound test rated material to a 180-pound test rated material with acceptance pull tests being increased to 20 pounds for the entire system.

This anomaly is closed.

14.4.3 Intermittent Lock of Universal Handling Tool In Suprathermal Ion Detector Fitting

While carrying the suprathermal ion detector experiment from the subpallet to the emplacement site, the experiment fell off the universal handling tool at least twice. The experiment sustained no visible damage and has been operating satisfactorily.

The universal handling tool fitting on this experiment is in the highest location above the lunar surface of any of the fittings and presents an awkward position of the tool for insertion, locking, and maintaining lock in the fitting ( fig. 14-46).

Corrective action includes training procedures to avoid inadvertent tool-release triggering because of the position of the tool. There are no present plans for the suprathermal ion detector experiment to be carried on future missions, and no other scheduled experiments have a similarly located fitting.

This anomaly is closed


14-5.1 Television Control Unit Clutch Slippage

During the second extravehicular activity, the camera could not be elevated as the unit approached the upper or lower limits of angular travel. The condition further deteriorated during the third extravehicular activity.

Elevation control is provided to the camera cradle through a friction clutch ( fig. 14-47) which allows manual override of the ground-commanded camera positioning. The camera-cradle pivot point is approximately 3 inches below the center of gravity of the cradle with the camera mounted. As the camera moves away from the horizontal position, the unbalanced moment becomes progressively greater, and a higher torque load must be supported by the clutch mechanism.

The elastomer clutch-facing material provided the required stable friction properties in the specification and qualification test temperature range (122' F, maximum). However, the maximum temperature on the television control unit during the third extravehicular activity has been calculated as approximately 180' F. Materials specifications show that the compressive strength of the elastomer degrades rapidly at this temperature, and ground tests with flight unit 4 verify severely degraded performance with time at elevated temperature.

The clutch is being changed to a metal-to-metal spring ring design in place of the elastomer disc. The clutch torque for Apollo 15 was set at 16 inch-pounds for ease of manual adjustment. For greater stability on Apollo 16, the new clutch is being built with a torque of 30 inch-pounds, which is still comfortable for manual positioning and is within design limits of the system, including the gear train (35 inch-pounds).

This anomaly is closed.

14-5.2 Lunar Communications Relay Unit Downlink Signal Lost

The lunar communications relay unit downlink signal was lost about 40 hours after lunar module ascent. The unit operated on internal battery power during the extravehicular traverses. Near the end of the third extravehicular activity, it was manually switched to lunar roving vehicle power in preparation for viewing ascent and for continuing television observations. The power distribution from the lunar roving vehicle to the television system is shown in figure 14-48. The lunar communications relay unit transmitter and television camera had been commanded on from the ground 13 minutes rior to the RF downlink- signal loss. The lunar communications relay unit status subcarrier had been commanded on 7 minutes prior to signal loss. The television camera was stationary and a 1-second incremental iris movement was occurring at the time of signal loss.

The flight data ( fig. 14-49) shows that the automatic gain control measurement began to fall followed by the video signal decay. This was followed by the decay of the lunar communications relay unit temperature measurement. The RF signal level then decreased below the ground receiver is threshold as indicated by complete signal loss. The overall loss of the downlink signal within 5 milliseconds is indicative of 28-volt d-c power loss. Decay of the temperature measurement is indicative of 16.5 volt d-c power loss. The lunar communications relay unit dc-to-dc converter ( fig. 14-48) supplies both the 28-volt and 16.5-volt d-c power. To verify loss of 16.5-volt power, an uplink voice signal was transmitted to key the VHF transmitter on. No signal was received on the Stanford 150-foot VHF antenna which indicates that the VHF transmitter, powered from 16.5 volts dc, was inoperative.

In laboratory tests, the fault which duplicated the flight data was the opening of the lunar roving vehicle power line prior to the 440-microfarad capacitor (figs. 14-48 and 14-49). The tests show that the decay time of the lunar communications relay unit 28-volt and 14-volt power is increased by discharging the 440-microfarad capacitor. Other induced faults resulted in shorter power decay times, affecting the received signal accordingly. The temperature measurement output (see thermistor in fig. 14-48) is proportional to the decay in 14-volt powcr. Consequently, the 6-percent decay of the flight temperature measurement corresponds to a 1.4-volt decay. This characteristic was duplicated when the lunar roving vehicle power line was opened. The 28-volt power decayed to 21.8 volts dc as the 14-volt power decayed to 12.6 volts. The RF transmitter power at this voltage will be decreased by 6.4 dB, and accounts for the total signal loss at this time since the ground receiver would be below its operating threshold.

The lunar roving vehicle power line has a 7.5-ampere circuit breaker forward of the 440-microfarad capacitor (fig. 14-48). Testing a 7.5-ampere circuit breaker under elevated temperatures (1800 F) and at vacuum conditions showed that the current capacity is also dependent on the connecting wire size because the wire provides a heat sink to the circuit breaker thermal element. The rover 7.5-ampere breaker used 20-gage connecting wire. Test results show that the breaker, with 20-gage connecting wire, at elevated temperatures and under vacuum conditions, will trip at 3.3 amperes. This corresponds to the calculated lunar communication relay unit load at the time of power failure.

A 10-ampere circuit breaker instead of the 7.5-ampere breaker and, in addition, a manual switch in the lunar rover circuit to override the circuit breaker after completion of vehicle activity are being provided for Apollo 16. Also, the lunar communications relay unit is being modified so that its internal 7.5-ampere circuit breaker is bypassed when operating in the external power mode.

This anomaly is closed.

14.5.3 Lunar Surface 16-mm Camera Magazines Jammed

The crew experienced film jams with the lunar surface 16-mm camera film magazines. Five out of eight magazines transferred to the lunar surface jammed, two were not used, and one successfully transported the film to completion.
Analysis of the returned magazines indicated two factors contributing to jamming.

Hardware analysis, air-to-ground voice tapes, and crew debriefing indicate that the lunar surface camera functioned properly, and the jammed magazines resulted from procedural errors. Corrective actions are to insure adequate crew training through scheduled prelaunch briefings, stress malfunction procedures and corrective actions, and put a removal flag on the tape.

This anomaly is closed.

14-5.4 Lunar Module Pilot's 70-mm Camera Film Advance Stopped

Near the end of the second extravehicular activity, the 70-mm camera ceased to advance film. The crew reported that the camera was again operational after return to the lunar module. The camera was used again on the third extravehicular activity; however, after a short series of exposures had been made, the failure recurred. The camera was used for additional photography during the transearth phase without recurrence of the problem.

Postflight analysis of the hardware included operational testing, disassembly and inspection, and measurement of battery charge. Operational testing with film loads indicated proper film advancement until approximately 200 cycles had been accumulated, at which time the failure mode was duplicated several times in succession. The film did not advance, although the motor was running. Disassembly and examination of the drive mechanism showed that the two set screws in the drive pinion were slipping on the motor shaft. After the last use of the camera during the mission, the crew had difficulty removing the magazine. This was caused by a rivet which had become detached from the camera magazine latch mechanism.

Corrective action is as follows: Flats will be ground on the motor shaft. A locking compound will be applied to the set screws when they are properly torqued against the flats. Also, epoxy will be applied to the tops of the screws to prevent them from backing off.

This anomaly is closed.

14-5.5 Difficult to Obtain Water From Insuit Drinking Device

After satisfactory operation during the first extravehicular activity, the mouthpiece of the insuit drinking device was displaced and the Commander was not able to obtain water during the second extravehicular activity. The Lunar Module Pilot was not able to actuate the drink valve of the insuit drinking device during either the first or second extravehicular activities.

After each extravehicular activity, the insuit drinking device was removed from the suit and all of the water consumed, thus verifying proper operation of the insuit drinking device drink valve. The problem was associated with the positioning of the insuit drinking device within the suit.

Ground tests using suited subjects and other equipment configurations indicated that the existing equipment provides the optimum configuration. The tests also showed that personal experience is essential to obtaining optimum individual positioning. Crew training is to include more crew experience in making the position adjustments required for the individual's needs.

This anomaly is closed.

14-5.6 Lunar Module Pilot Oxygen Purge System Antenna Was Damaged

The crew reported that the Lunar Module Pilot's oxygen purge system antenna was broken off near the bottom during communications checkout prior to the second extravehicular activity. Previously, a notch had been observed in the antenna blade(see fig. 14-51).

Antennas broken in training have shown similar flexure breaks. Observation of the notch edges of the returned antenna indicates that the notch started as a partial break in flexure, followed by material being torn out the rest of the way. Test results of the returned antenna indicated that the physical properties of the blade material were satisfactory with no excessive brittleness.

A flap which covers the entire antenna will be added for Apollo 16 to protect the antenna while the oxygen purge system is stowed and during unsuiting after extravehicular activities. The antenna will not be deployed until after egress to prevent it from being damaged inside the cabin or during egress.

This anomaly is closed.

14-5.7 Retractable Tether Failure

Both retractable tethers failed during lunar surface operations; the Commander's tether cord broke during the first extravehicular activity, and the tool clamp came off the end of the Lunar Module Pilot's tether. The Commander carried the standard 3/8-pound pull tether which consists of a case, a negator spring wound reel-to-reel on two spools, and a 30-pound cord wound on a spool mounted to one of the spring spools ( fig. 14-52).

A tool clamp is attached to the external end of the cord. The Lunar Module Pilot carried the optional, somewhat larger, 1-pound pull tether of the same design.

Disassembly of the Commander's tether showed that the spring had expanded off the spool, snarled, and jammed against the case as the result of a no-load release of a slack cord (fig. 14-53).

The cord had broken against a sharp edge of the spring when an attempt was made to extend the tether after the jam. The failure mode with the release of the slack cord is repeatable. Disassembly of the Lunar Module Pilot's tether showed that both the bowline and the figure-eight knot attaching the cord to the clamp had untied ( fig. 14-53) and this allowed the cord to retract into the housing. Changing this knot to an improved clinch knot will provide a more secure and permanent attachment. Crew training will emphasize proper use of the tethers.

This anomaly is closed.


14.6.1 Deployment Saddle Difficult To Release From Vehicle

The lunar roving vehicle deployment saddle was difficult to release from the vehicle during the final stage of deployment operations.

The causes of this problem are twofold and interrelated.

Ground tests have shown that if the partially deployed lunar roving vehicle resting on the surface but not yet detached from the saddle and lunar module, is rolled either to the left or right, the saddle/rover chassis interface will bind. The interface can be released, and the saddle dropped to the ground by one crewman adjusting the roll back to zero while the other taps the saddle with a hand tool. The corrective action is to insure adequate crew training.

This anomaly is closed.

14.6.2 Volt/Ammeter Inoperative

The lunar roving vehicle battery 2 volt/ammeter was inoperative upon vehicle activation, and remained inoperative throughout the traverses. Problems with the meter were experienced during its initial development; however, after a more rigid acceptance test program was implemented, the earlier problems were cleared. The flight problem was not duplicated during any of the ground tests. Since the instrument is not essential for the operation of the vehicle, no further action is being taken.

This anomaly is closed.

14.6.3 Front steering System Inoperative

During initial lunar roving vehicle activation, the front steering was inoperative. Electrical checks were made which verified that electrical power was being supplied to the front steering system. Unsuccessful attempts were made to manually rotate the wheels about their steering axes and to detect steering motor stall current on the ammeter. The forward steering circuit breaker and switch were cycled without any apparent effect. Consequently, the front steering was switched off for the first traverse. During preparations for the second traverse, the forward steering circuit breaker and switch were cycled and front steering was operative; however, the time that front steering capability was restored is unknown. Front-wheel wandering did not occur during the first traverse, indicating a mechanical problem. The steering continued to function properly for the second and third traverses. During the second traverse, the rear steering was turned off temporarily and wandering of the rear wheels occurred.

The most likely cause of this anomaly is motor and/or gear train binding, as indicated by the inability to drive back through the linkage and gear train by manually pushing against the wheels. Electrical causes are possible, but less likely.

The front steering system of the Apollo 16 lunar roving vehicle is currently being analyzed because of an intermittent failure of a similar nature. Manually pushing against the wheels would not always drive back through the linkage and gear train and the motor stalled at limit current for 0.8 second during a test of this condition.

This anomaly is open.

14.6.4 Lunar Roving Vehicle Seat Belt Problems

The following seat belt problems were experienced throughout all traverses

The main causes of these problems, in addition to insufficient belt length, were insufficient belt rigidity and lack of visibility of the securing operation.

New, stiffer seatbelts with an over-center tightening mechanism will be provided for Apollo 16 to eliminate adjustment after each ingress and to provide more tightening capability.

This anomaly is closed.