MEMORANDUM for Acting Chief, Research Division
FROM: X-15 Research Project Office
SUBJECT: Research Report on flight 3-39-62
I. Purpose of flight
A. Flow field experiment - the flight requirements were:
| Data point | a (±2°) | V (fps) |
| 1 | 10° | 5000 |
| 2 | 17° | 5000 |
| 3 | 10° | 4700 |
| 4 | 17° | 4000 |
| 5 | 10° | 3500 |
B. Surface discontinuity heat transfer test.
C. Skin friction.
D. Boundary-Layer Noise.
E. Nose Gear modification checkout.
No specific flight requirements were made for these remaining experiments.
The flight plan was based on the flow field experiment flight requirements. Flight was flown by M. 0. Thompson.
The first two data points were obtained prior to burnout. In attempting to obtain the third data point, the pilot experienced violent oscillations in all three axes. These oscillations shall be discussed under section IV. The fourth data point was obtained satisfactorily. The fifth data point was initiated, but rapidly changing conditions (H and q) and energy management considerations, forced a reduction of the duration of this point. Table I indicates the degree of success reached in obtaining the requested points.
TABLE I
| Data
Point |
alpha
degrees |
Velocity
fps |
Mach
No. |
Duration
secs |
q
psf |
| 1 | 10° +2/-1 | 4330-4900 | 4.43-4.99 | 10 | 704-692 |
| 2 | 17° +0.8/-0 | 5120-5340 | 5.18-5.39 | 12 | 624-568 |
| 3 | The oscillations occurred at this time period. | ||||
| 4 | 17° +0.6/-1.5 | 4960-4300 | 5.04-4.44 | 14 | 534-610 |
| 5 | 10° +0.1/-0.2 | 3910-3730 | 4.03-3.87 | 5 | 698-770 |
There were no extended periods of essentially constant q during this flight except during the scheduled data points. (0 - 10 psf/sec is considered constant). However, q rate remained relatively constant (» 12.5 psf/sec) over a long period of time (» 40 sec).
A peak velocity of 5420 fps was reached instead of 5100 fps due to an extended burning time of 8.5 sec on the engine. The pilot did not sense that the engine did not burn out due to the minimum thrust setting, the small longitudinal acceleration, and the high normal acceleration.
The flight was flown exceptionally well under the existing circumstances.
III. Instrumentation by experiments
Flow field; 35 channels on oscillograph 0-7, 0-23, manometer P-2 and P-8.
Skin friction; 23 channels on oscillograph 0-23, manometer P-27 and P-11.
Ablative; 12 channels on T.C. oscillograph 0-25.
Boundary-Layer Noise; 4 tracks on tape recorder 38.
Nose Gear checkout; 5 channels on oscillograph 0-7.
B. Results
2. Skin friction: The skin friction gage recording on oscillograph 0-23 operated satisfactorily. All recorded data appears to be acceptable.
3. Heating tests: A small-wave roughness panel was installed. No malfunctions were noted and temperature data appears satisfactory.
4. Nose gear modification checkout: No instrumentation malfunctions were noted on the flight. All recorded data appears to be acceptable. However, the following is quoted from a memorandum dated January 26, 1965, from the Structural Mechanics Section, "Scoop hook CPT data is not included ( in the memo) as this data is questionable due to shortcomings in CPT location."
The following table reviews up-lock hook loads:
| Flight | Time after launch sec | Up-lock hook load |
| 3-37 | 135 | 3468# |
| 3-38 | 133 | 3250# |
| 3-39 | 134 | 3500# |
Nose gear modification checkout instrumentation has been installed for the last four flights and the study appears to be far from completed. The only temperature available for correlation is obtained from a thermocouple at station 72 on bottom centerline. This position is just behind the nose gear door. On flight 3-37, for example, a peak up-lock hook load of 3468# was measured at 135 sec and probably results from a maximum temperature gradient through the door. However, peak temperature of 981° at station 72 was reached at 269 sec. It does appear that one could obtain correlation between loads and differential temperatures on the door itself, and consideration must be given toward implementation of this research. Temperature measurements around the fuselage in the vicinity of the nose-wheel also may yield information on diametrical expansion.
In the opinion of this office, the research knowledge to be gained should prove valuable and unless good correlation with various temperatures is obtained on the gear problem, the study will be considered far from complete.
5. Boundary-layer noise: Malfunctions occurring in the instrumentation still are not being noted on the standard "squawk" sheet.
Very recently, it has been brought to the writer's attention that one of the pressures recorded on the BLN rake (P08101) has had difficulties for the last 3 flights. Apparently the difficulty cannot be found on pre- and post- flights and occurs only from launch. It is anticipated that this problem shall be solved prior to the next flight.
The engine was removed to provide access to the main landing gear. An engine run was accomplished on January 8, 1965. An APU was replaced and a satisfactory run was made on January 9, 1965.
The aircraft was weighed on January 11, 1965 . Basic weight was determined to be 13,940 pounds.
Mating was accomplished on January 11,
Flight was not attempted on January 12, 1965, due to Air Force support aircraft problems.
The aircraft was flown on January 13, 1965.
B. Stability Control Problem - Dutch Roll Oscillations
As per flight plan, the pilot obtained data point #2, while in a left bank.
The third data point was to be obtained after burnout and after establishing a 60 ° right bank.
About two seconds before burnout the pilot reversed from the left bank (f = 86.5°). As the airplane rolled level, engine burnout occurred. The roll was continued into an 83° right bank at an angle of attack near 5°. The pilot attempted to stabilize the right roll maneuver and to rotate to an angle of attack of 10°. Internal recordings indicate that pitch and lateral-directional oscillations occurred during the following 10 second interval. Surface and servo limiting was indicated on both damper servos and position limiting of the starboard servo occurred intermittently throughout this interval.
The adaptive gain in pitch of the MH-96 system was at maximum during these oscillations. The pitch rate and frequency during the oscillations were sufficient, with the system at maximum gain, to produce servo and surface rate limiting. The lateral oscillations experienced by the pilot result from the loss of aircraft (mechanical) damping in roll during intervals when the surfaces were rate limited in pitch. Internal recordings and analytical computations indicate the MH-96 adaptive control system functioned in a normal manner.
Flight conditions, which the pilot could expect to encounter, had been evaluated on the X-15 simulator prior to the flight. Servo rate limiting and lateral directional oscillations were not indicated during the simulator study. Attempts to duplicate the oscillations immediately after the flight were unsuccessful.
In an attempt to provide a solution to this problem, two
areas of investigation will be conducted through a study contract with
Honeywell in association with NASA personnel. The MH-96 adaptive system
will be evaluated to determine methods of restricting servo motion to provide
pitch and roll damping under similar conditions. To improve the dynamics
of the X-15 simulator to agree with the aircraft control system dynamics,
the X-15 simulator will be improved through the addition of hydraulic pumps
with higher flow rates, reduction in hysteresis, rerigging of the mechanical
system, and analysis of the adaptive system operation to provide a conservative
estimate of X-15 flight with this system. The simulator has been re-rigged
by an X-15 crew and hydraulic pumps have been acquired. Analysis by NASA
personnel and Honeywell representatives will continue.
Aerospace Engineer