B. The X-15 pilot's regulated O2 supply warning light flickered during flight l-A-78 and corrective action was not reported in the previous report. The post flight investigation revealed the oxygen regulator provided a nominal 68 psig pressure, which is at the low end of the tolerance. A regulator with an 80 psig setting and a new pressure switch were installed, and satisfactory operation was obtained during flight 1-50-79.

C. As a result of the unscheduled extension of the micrometerorite experiment during flight 1-50-79, a re-evaluation of the control circuitry was made by NAA and NASA engineering; wiring configuration was also verified. No single malfunction could be found to explain the in-flight operation. A revised operational philosophy was advanced to replace automatic sequencing with pilot initiation and termination inputs. NAA was requested to evaluate a door open .limitation for flight. If conditions with the micrometerorite door open are prohibitive, a redesign will be required to provide aerodynamic door closing to compensate for the singularity of the drive motor. The motor must operate to obtain a closed door condition in the present configuration. The experiment is inactive during the ensuing investigation, and the door is secured with metal clips and screw fasteners.

D. Contact was made between the speed brakes inner skin and the MIT tail cone box forward face during the previous flight. The flight plan required 35° extension of the speed brake at 28 seconds burning time. Ambient temperature clearances were .070 inches minimum and trimming was accomplished to increase clearance to approximately 0.120 inches.

E. A temperature survey of the wing outer panel and leading edge was considered to determine the thermodynamic effects of the wing pod installation. However, a review of the wing T.C. history indicated extensive deterioration had occurred requiring an extended effort to locate and checkout each T.C. Therefore the decision was made to use thermo-paint in lieu of the T.C. survey. One T.C. was installed on the inner surface of the LH pod skin one inch forward of the wing/pod apex. A T.C. was available from the micrometerorite experiment.

F. Replacement of the IFDS computer (Verdan) S/N F-1 was accomplished after flight 1-50-79 because of a characteristic of reverting to the "halt mode." Computer S/N F-2 demonstrated the "halt mode" tendency only once during post installation ground operation. The subsequent preflight activities were normal, and a consistent compute mode was obtained.

G. The engine second stage pressure cockpit display system was calibrated satisfactorily. A frozen pressure sensing line is suspected as the cause for slow response on the prelaunch engine procedure of the previous flight.

B. Relocation of several pieces of instrumentation was accomplished for better access or protection of fragile parts. The FM NEFF power supply and IFDS J-box positions were exchanged, and the FM/FM J-box and TATP 300 chassis were exchanged.

C. Switches were added to the cockpit center pedestal to control power to the air density, sky brightness, and MIT horizon scanner. These switches allow reduction of system power-on time and reduction of electrical load in-flight in case of loss of one alternator or APU (MIT heater power is 800 watts).

D. The MIT horizon scanner was installed in the tail cone box. The configuration included a fixed gimbal and a VDR4 camera. IFDS vertical reference was not utilized.

E. A lanyard was added to the pilots physiological seat-to-suit disconnect to obtain proper separation. Hi-Johnson multistrand control cable, with an approximate 100 lb. pull test, was used. The attachment to the plug shell is provided by a drilled hole in a raised ring on the shell.

F. Instrumentation was added to the nose gear to investigate motions of the scoop door latching hook and roller. Rotation of the hook is measured by a reel type CPT, and a deflectometer (spring steel fingers with strain gage) measures relative movement between the scoop door roller and uplock hook. The effects of fuselage diametrical and longitudinal expansion on hook roller engagement are expected to be defined for the "old" nose gear configuration.

G. A high range absolute pressure transducer was added to the airspeed recorder to measure ball nose total pressure.

H. The air density experiment (Univ. of Mich.) was operative including a T.C. to monitor the environment of the radioactive source.

I. Although the micrometerorite experiment was inoperative, a clean capsule was installed to measure contamination in the closed system. One T.C. from the micrometerorite experiment was used to monitor the skin inner surface temperature at the pod/wing apex.

J. The nose BCS F-3 and F-4 panels were prepared with ablative materials. Langley Purple Blend was installed on F-4 and Martin 255 on F-3.

K. The Sky brightness experiment (Nortronics) was not installed for flight because the experiment is concerned with higher altitudes. Ballast was installed.

B. Rework of X-15-3 nose-gear uplock and scoop door mechanism produced an interference between gear mechanism and structure. Inspection of X-15-1 was accomplished, but interference was not evident with the unmodified configuration. The nose-gear uplock to roller clearance was measured at .040 - ,050 inches and readjusted to specification .010 - .020 inches.

C. APU load check runs were satisfactorily accomplished on November 4, 1964.

D. The engine vibration system (engine S/N 107) was replaced because of failure of the "B" unit to meet output specifications.

E. A cockpit leak check was made after repair of cable glands with a leakage rate of 70 cfm.

F. Delays of X-15-2 caused by weather required rescheduling of X-15-1 for 11/12/64 and 11/13/64, 11/17/64 and 11/18/64. Because of snow, a repeat of aircraft preflights was initiated on 11/17/64 for a flight scheduled on 11/30/64.

G. Because of experience on X-15-2 during a captive flight gear check, where slow nose gear operation was obtained, an inspection of the nose gear down snubber was accomplished on X-15-1. Snubber action was demonstrated using a 100 lb. weight and timing the extension cycle; five out of six units tested including the unit removed from X-15-1 demonstrated an 8-9 second cycle time, the remaining unit required 12 seconds. The snubber was replaced because it did not exhibit smooth operation when operated by hand.

H. The wing pods and outer wing panels were painted with thermopaint to evaluate the aerodynamic effects of the pods. The dark green (1100° F range) was used on the left hand side and the light green on the right hand side.

I. APU load check runs were again accomplished on 11/19/64.

J. A realignment of SAS was necessary because of discrepancies in the preflight. No problems were encountered during realignment.

K. On final closeout checks, a pinhole leak was found in the elevator lid seal requiring replacement. A problem of fitting the new seal to this elevator lid was encountered; survey of the seals in stock indicated poor fit, with some seals as much as three inches too short in circumferential length. A verification cockpit leak-check was made at 72 cfm after seal replacement.

L. The aircraft was mated to B-52 #003 on 11/25/64 and held in backup status for the flight demonstration on 11/30/64.

M. Aircraft servicing through APU H202 was accomplished on 12/1/64 before weather required cancellation. An IFDS run was accomplished to reload the flight profile with a Delamar launch (formerly scheduled Mud Lake was not acceptable because of snow and ice).

N. Excessive winds prevented a flight on 12/2/64.

O. The aircraft was completely serviced, and flight preparations completed up to the B-52 ground to alternator power transfer on 12/3/64; results discussed in part IA. The flight was canceled and tests conducted to provide an acceptable electric power transfer procedure to prevent loss of the IFDS.

P. Flight l-A-80 was accomplished on 12/4/64 with Capt. Joe Engle pilot and B-52 #003 carrier aircraft. Leakage of the fuel vent valve occurred during the B-52 takeoff roll. Because of the unknown quantity of fuel lost, the flight was canceled and propellant jettison accomplished at 25,000 feet. The IFDS was placed in altitude inertial mode and performed satisfactorily.

Q. After flight, l-A-80 a series of system operation verification tests were made including a 2 minute APU run with operation of generator reset, full-throw control cycle, flap and speed brake cycle, and ball nose turn on. No discrepancies were noted. The APU run was accomplished to allow extension of the aircraft system preflights which were out of date as of 12/4/64.

R. The gas leakage through the fuel vent-valve manual selector was checked satisfactorily and the fuel vent-valve replaced to correct the vent-valve malfunction of flight l-A-80.

S. Weather caused delay of X-15-3 on 12/7/64 and 12/8/64, resulting in delay of X-15-1.

T. An IFDS reload and alignment confidence run, and the engine lube-oil, hydraulics, and electro-mechanical operation checks were accomplished to complete flight preparations for 12/10/64.

B. The No. 1 APU tank pressure rose slowly after source service to approximately 60 psi at APU start. Leakage is occurring through the APU helium shutoff valve.

C. An exceptionally high cabin helium source pressure of 3800 psi was obtained after servicing, and a resultant relief cycle reduced the pressure to 3600 psi. Warming of the source bottle by external IFDS cooling after an extended LN2 bay cold soak is the suspected cause. Subsequent leakage of cabin source was noted during the outbound track with approximately 1500 psi remaining after landing.

D. The ground to B-52 alternator power transfer was successfully accomplished without loss of IFDS, employing the procedure of setting the right-hand aft alternator first. General discussion is given in Part lA of this report.

E. The pilot reported loss of the pitch SAS ten seconds after launch and a reset was accomplished. Judging from the aircraft stability characteristics in pitch throughout the flight, the pitch did not perform normally although the channel light was not illuminated. Also, the expected disengagement of the pitch channel did not occur on touchdown with the landing SAS disengage servo lockup. After flight, hangar tests were conducted with the gyro assembly on the rate table. First tests indicated rates imposed on the gyro caused neither movements of the horizontal stabilizer or normal trip outs. However, with subsequent testing satisfactory operation was obtained with no malfunction. Further tests are planned to pursue the problem.

F. The IFDS operated throughout the flight with excellent performance as evidenced by post landing readouts.
 
 
 
 

2. Install redesigned CEU to computer harness incorporating computer loader outputs. Compact wire dressing techniques were employed to reduce congestion in front of the computer. Requirements are met for access to computer inputs for trouble shooting and computer programming.

3. Relocation of the IFDS Interface "J" Box. Access to the original "J" box location was marginal. The F/M NEFF power supply and IFDS "J" box positions were exchanged without rewiring.

4. Power Control Configuration. Operation of the IFDS has resulted in an inconsistent shutdown condition of the computer preparatory for flight turn-on. With "compute" mode set up during hangar preflight, the computer has been found in the "halt" mode, requiring reloading before preflight alignment.

2. Experiment Power Control Switches. Extended periods of A/C power is compromising the electronics of the sky brightness, air density experiments; a provision for power control switches in the cockpit allows limitation of operating time. Power control for the MIT Scanner is required in case of APU/Gen failure where reduction of electric load is desirable.

3. W/O 4031 will be accomplished to add a high pressure cell to the airspeed recorder total pressure configuration.

2. Return the Micrometerorite package (LH forward pod assembly) to NAA, LAD, for investigation of unscheduled extension during flight 1-50-79. Aircraft wiring will be thoroughly checked against the print and a functional demonstration of the aircraft system relays will be made.

3. Complete replacement of deficient cherry rivets in the lower window fairing structure with proper Jo bolts.

4. Investigate cause for high #2 mixing chamber temperature during flight, including cause for the #2 N2 control C/B disengagement during flight.

5. Calibrate the stage 2 igniter pressure system (pilots display).

6. Alter the pilots alpha indicator by extending the graduation marks for 10 and 20 degrees.

7. Relieve structural interference between IFDS computer, AGE harnesses, and shelf supports at RH, outboard, aft corner of IFDS computer.

8. Make an effort to improve cockpit leakage rate wherever possible. Previous preflight and postflight rates were 80 CFM.