Subject: Flight Suitability of the Number One X-15
 
 

The Ad Hoc Committee respectfully submits the attached report on the flight suitability of the number one X-15. The committee concurs with the underlying philosophy of the wiring configuration changes. With satisfactory demonstration of system performance during planned ground operations the vehicle, in the committee's opinion, will be suitable for flight.

For the Committee
 
 
 
 

Perry V. Row

Committee Chairman
 
 

Background

During Flight 1-73-126, on June 29, 1967, an electrical system failure shut the engine down at 69 seconds of burn time and all electrical systems ceased functioning. Since the hydraulic gages are electrically powered, the pilot had no indication of whether both electrical and hydraulic power or just electrical power was lost. Attempts to restart the APU's were unsuccessful until switching the emergency battery on, after which the No. 1 APU was started. Landing was made at Mud Lake with hydraulic power but without generator power.

Post-flight investigations, conducted after the airplane was returned to Edwards, did not immediately show evidence of any problem area. After a thorough inspection and functional checkout, a complete APU ground run was made with normal operation. System checks of the engine, SAS inertial system and MIT experiment were made with intentional shutdown of one APU during engine-run time. Operation was normal and No. 1 generator carried the load when No. 2 generator was switched off.

During normal operation, the APU control system provides a constant speed drive regulating AC generator output to 400 ± 4 Hertz (Hz). The voltage regulator provides control of voltage to 115 ± 1.5 volts. Each AC generator system contains a protective circuit for automatic control of an over-voltage or an under-frequency condition. These conditions are 135 volts and 370 Hz respectively. In the case of a generator failure, relays disconnect the particular AC buss and reconnect the failed buss to the opposite AC system. The emergency battery provides for control of the AC system during the initial engagement process. The X-15 electrical system employs components and control circuits which are comparable to all aircraft electrical systems.

Any electrical malfunction in the aircraft which can create an overload which lowers the generators frequency to the 370 Hz level, or any loss of the electrical signals to the various relays which provide the circuits for controlling the generator, would result in loss of that generator and transfer the load of one AC buss to the opposite generator. Components which are suspect to the incident of this report would include shorted or grounded wires and connections, relays, circuit breakers, as well as the circuits internal to the generators. It can be theorized that the malfunction causing the loss of the first generator could be transferred to the second generator which also bears the additional electrical load, unless the malfunction was in the APU drive. This would only result in the automatic transfer of the load of both generators to the remaining operative circuit.

The problem of generator overload is not new to the X-15, and considerable testing was accomplished in the past to qualify the X-15 generator to an approximate 9 KW limit. At the same time, the concept of an automatic secondary buss disengagement of noncritical loads was tailored to fit the single AC generator operation condition within the 9 KW limit.

A flight data survey of both T/M (up to first electrical failure) and internal data (seven seconds longer until complete electrical failure) indicated nothing unusual until the first failure. Apparently the No. 2 generator failed with a clean loss of output to initiate the ensuing loss of electrical power. There was no indication of No. 2 APU failure, since T/M was lost abruptly at this point, and there were no APU parameters on internal data. Likewise, there was no indication that APU No. 2 did not shutdown. The No. 2 generator currents, which were recorded internally, dropped to zero and there was a slight increase in No. 1 generator currents indicating that it had picked up at least a portion of the load. The actual increase in load on No. 1 generator could not be determined since full scale output of the monitoring devices was just slightly above normal load for the generator. The internal recording of No. 1 Af (A-phase of 3-phase generator output) voltage indicated an intermittent voltage signal similar to that caused by turning the generator field exciter on and off rapidly for the seven seconds prior to complete electrical failure.

During the week of July 21 the Air Force Western Test Range (WTR) launch monitoring experiment was subjected to environmental testing in the altitude chamber. No discrepancies were observed for altitudes less than 45,000 feet and temperatures greater than 0°F. A functional check at 100,000 feet and -60°F disclosed large power transients and the frequency converter fuse was blown. The fuse was replaced and another run under the same conditions was attempted. Power transients between 90,000 and 95,000 feet resulted in the blowing of two 15-ampere fast-blow fuses in the laboratory wall power. The experiment would no longer function properly after this run. Investigation showed evidence of arcing at the experiment Af and Cf pins of the power plug. The pins were burned off and the plug insulation was split between all three phase pins indicating strong arcing between the A phase and the other two phases.

Although it is thought that the WTR experiment plug failure precipitated the in-flight electrical failure, a complete failure analysis was impossible because the flight failure could not be duplicated (possibly because the environment could not be duplicated) during ground investigations. Sufficient evidence of wiring discrepancies which could contribute, through concurrent malfunctioning, to the flight failure were found to justify rewiring of the aircraft. Originally aircraft electrical wire routing of comparable APU and generator circuits was not divided into separate wire bundles to provide protection against the case of a single malfunction affecting both systems. A selected rewiring effort was accomplished on the aircraft due to the fact that many isolated wiring discrepancies could have contributed to intermittent conditions related to a portion of the flight experience but could not be used to explain the whole problem. Several other electrical configuration changes were made, as follows, and for which this Ad Hoc Committee has been commissioned to evaluate.

Configuration Changes

The WTR experiment and MIT experiments have been added to the secondary buss which drops out in the event of one generator failure. Previously this buss carried the No. 2 mixing chamber blower, the right hand windshield heater and the secondary buss relay for a total load of approximately 1.5 kw. Now, with the addition of these experiments, the secondary buss carries approximately 2.9 kw (which drops out when one generator fails).

The flap motors, third-skid, stick pusher, and hydraulic gages (discussed later) have been added to the emergency battery buss. The Sonotone (Part No. Z-19876) 24-volt battery has been replaced with a higher capacity Sonotone (MA-300H) 24-volt battery (used in F-104 aircraft).

The WTR experiment power plug was formerly wired from two busses. It now obtains power from one buss, thus permitting greater distance separation of the high potential pins. This plug has also been potted. The experiment circuit breakers have been reduced from 10 amperes rating per phase to 5 amperes rating per phase.

The routing of APU control wiring has been changed to provide better physical separation of the No. 1 and No. 2 systems in order to reduce the possibility of a single electrical problem shutting down both APU's. All switches and circuit breakers have been replaced with new components. The power panel has been redesigned (similar to X-15-3) to provide better routing, less bulky wire bundles through the panel-box and greater clearance between busses.

With the loss of electrical power the hydraulic pressure gages were formerly inoperative, thus the pilot had no indication of whether the APU's were shutdown completely or whether just the alternators were inoperative. A dual indicating DC-powered gage, operating off the emergency battery, has been installed which will provide the hydraulic pressure indication with generators off. The gage was environmentally checked to 100,000 feet altitude and at -65°F temperature with no appreciable loss of accuracy.

The overvoltage relay in the AC generator protective units and the frequency control relay in the generator system have been revised to control the generator exciter windings rather than the generator field windings. This change was made at NAR's request and Engineering Order to provide less severe load conditions on the protective relay contacts.

Functional check runs

A series of runs were accomplished on December 21, 1967 for two purposes. The first two were to check and adjust the voltage regulators. The second two runs were to obtain combined operation data.

These runs were generally successful except for two discrepancies. When an attempt was made to retract the PMR experiment the mirror would swing into the "rest" position and then swing away and the experiment would not retract. The No. 1 generator was then turned off, which turns off PMR power through the secondary buss. The PMR experiment then retracted using the lanyard to stow the mirror (emergency type retraction). The rest of the run was accomplished as per checklist. Post-run checks in the hangar revealed no cause for this discrepancy. Subsequent operations, even with two separate power supplies simulating APU operation, was satisfactory. Also during the run the precision attitudes looked erratic on APU power, but were good on ground power. After the run, the transformer-rectifiers (TR's) were removed, checked, and replaced. A hangar check with two separate power supplies revealed the apparent noise problem was not corrected. Since the TR's output was too noisy, a regulated power supply has been procured and will be installed at a later date after undergoing environmental testing (after the first flight).

Planned Future Operations

Before next flight, it is planned to (1) complete a normal preflight (2) make an engine run, (3) repeat the combined APU-BCS-Powered experiment run, (4) go through a mated simulated flight rehearsal and (5) have a captive flight. The simulated flight rehearsal will be an X-15/B-52 compatibility check as well as an instrumentation checkout of on-board data acquisition systems and the control room data readout system. All systems and active experiments will be operated except the BCS and YLR-99 engine. The captive flight will be a normal pre-launch procedure including engine operation through igniter idle.

Conclusions

A satisfactory explanation of how the complete electrical system was failed during Flight 1-73-126 has not been found. Conjectural explanations could not be proven, but were the basis for rewiring the aircraft in its present configuration. The committee agrees with the logic for making the configuration changes.

The planned operations prior to flight should provide adequate checkout of all systems, within the capability of simulating flight environments. Satisfactory demonstration of system performance through these operations will, in the committee's opinion, qualify the aircraft for flight.

Recommendations