NASA FRC

NASA FRC

October 26, 1967

MEMORANDUM for Assistant Chief, Research Projects

Subject: Preliminary Report on X-15 Flight 2-53-97

Summary

Flight 2-53-97 was flown by Major William J. Knight on October 3 1967 , for the purpose of obtaining data for:

a. Evaluation of the full-coat ablatives.

b. Stability and control evaluation with the dummy ramjet, external tanks, and ablatives.

c. Ramjet local flow measurement.

d. External tank separation characteristics.

e. Fluidic temperature probe.

The maximum velocity attained was 6650 fps (4534 mph and a Mach number of 6.72) at an altitude of 98,000 feet. The maximum altitude reached during the flight was 99,000 feet. The engine burn time at 100 percent thrust was 139.0 seconds.

The primary flight objectives were attained. The pilot considered the handling qualities of the aircraft to be acceptable. The external tank separation was satisfactory and the external tank recovery system worked properly, however, each external tank drogue chute slug impacted with the tank rupturing the skin. The external tanks are being repaired. The ablative material functioned satisfactorily with the exception of a complicated shock wave impingement on the ramjet pylon which produced high local aerodynamic heating. This problem is currently being analyzed. As a result of this local aerodynamic heating, the ramjet instrumentation failed approximately 25 seconds after shutdown and the dummy ramjet shape separated from the aircraft over the Edwards bombing range. Acceptable local flow and fluid temperature probe data were obtained during the prelaunch and boost phases of the flight.

Flight Track and Profile

The radar track and profile are shown in figure 1. The launch was from Mud Lake number 1 at an indicated altitude of 43,750 feet and a velocity of 225 KIAS.

The pilot reported shutting down the engine as planned at 6500 feet per second. The maximum velocity estimated from the cockpit film was 6630 feet per second. The total engine burn time was 141.4 seconds with 139.0 seconds at 100 percent thrust. Compensating for the thrust buildup and decay, the burn time compares with the simulator planned time of 141 seconds, however, the aircraft velocity was 130 feet per second higher than planned.

The time histories of aircraft weight and longitudinal center of gravity are shown in figures 3 and 4, respectively.

Stability and Control

A time history of stability and control parameters telemetered during the flight is shown in figure 2.

The pilot reported the longitudinal control task was more difficult than it appeared in the simulator because it was more sensitive. He indicated it was difficult to maintain a trimmed angle of attack within 2 degrees of that desired. The pilot noted the rudder effectiveness and directional stability appeared to be lower than predicted by the simulator. The overall handling qualities were still considered acceptable.

The pilot reported the launch was smooth (the maximum roll-off after launch was 14 degrees) and indicated the roll control was better than the last external tank flight. The roll angle during the external tank portion of the flight did not exceed ±8 degrees.

After launch, the pilot reported pulling up to an indicated 15 degrees angle of attack and experiencing buffet. An angle of attack of between 12 and 13 degrees was maintained during rotation. A maximum normal acceleration of 2.0 "g" and a maximum dynamic pressure of 570 psf occurred during the rotation. The planned 35 degrees pitch attitude was maintained within ±l degree during the climb.

The pilot felt the handling qualities of the aircraft with the external tanks were quite good to the maximum angle of attack of 15 degrees. During this flight, the pilot reported he was better able to control the left roll tendency caused by the heavier LOX tank . The external tanks were ejected 67.4 seconds after launch. The angle of attack peaked at 5 degrees, then dropped to -2 degrees during the trim change. After the tank ejection, the planned 2 degrees angle of attack was maintained within ±1 degree. After leveling at 99,000 feet, the pilot reported it was most difficult to hold a constant angle of attack of 6 degrees required to maintain a zero rate of climb. Internal data indicate the angle of attack varied between 5.5 and 7.5 degrees during this interval.

The stability and control maneuvers performed after shutdown are shown in table I. The pilot reported the first rudder pulse did not oscillate as expected and indicated the period was longer. After extending the speed brakes, another yaw pulse was performed which was described as deadbeat and not as expected from the simulation. Internal data indicates the damper was set at low rather than off as planned. The pitch pulse in low gain that followed was described as being like the simulator. After the last yaw damper disengagement, the pilot failed to re-engage the yaw damper and flew the remainder of the flight without yaw augmentation damping.

External Tanks

Aircraft operation with external tanks installed appeared to be satisfactory. The pilot switched to internal propellant flow at 61.7 seconds after launch. After switching to internal propellant flow, the pilot pushed over to 5 degrees angle of attack for the tank ejection. The tanks were ejected 67.4 seconds after launch. The conditions at tank ejection were:

Planned Actual

(Beatty radar)

Velocity, fps 2,000 2,083

Altitude, feet 66,000 70,930

Dynamic Pressure, psf 340 301

Angle of Attack, deg 5 4.2

Normal Acceleration, "g" 0.5 0.25

At tank ejection the angle of attack instantaneously increased to 5 degrees, then during the trim change, the angle of attack decreased to -2 degrees. The pilot reported the ejection transient to be harder than the last one experienced (flight 2-50). The maximum transients in normal acceleration were from 1.8 to -0.4 "g" following the external tank ejection.

The external tank recovery system appears to have performed satisfactorily, however, each tank drogue chute slug impacted with its tank rupturing the skin and the drogue chutes remained with the tanks. The tanks impacted in the Cactus Mountain Range as indicated on the map in figure 5. Photographic coverage of the tank separation and recovery system operation were obtained from the AEC Tonopah Test Range. The tanks were retrieved from the impact area and are repairable.

Thermal Protection System

The ablative coating which completely covered the airplane for this flight, had been flown on the previous flight and was refurbished for this flight. The performance of the ablative coating was considered good in all areas except the lower ventral stub. About 90 percent of the aircraft surface was intact except for a very thin layer of surface degradation. The wing and tail leading edges were charred to approximately half the depth of the Mach 7.4 design thickness. Measured temperatures reached a maximum of 34°F on the lower fuselage at the nose gear door and 240°F on the lower fuselage at station 340. The maximum leading edge temperature at the wing midspan was 110°F and at the same station, but back on the lower quarter-chord, the temperature reached 310°F. The ablative coating was designed to give a 600°F skin temperature on the M » 7.4 design mission.

The unprotected right-hand windshield on the canopy experienced redeposition of ablation products to the extent that the pilot's vision was obscured, as had been anticipated. (Adequate visibility for landing is being provided by the left-hand windshield which is mechanically covered until just prior to landing.) Areas of the aircraft surface affected by local high-heating rates due to protuberances may require localized use of higher density ablator. A total failure of the ablator and melting of the Inconel-X dummy ramjet pylon leading edge occurred because of the unexpected magnitude of shock wave boundary-layer interaction heating. Maximum skin temperature is estimated to have been in excess of 2700 degrees F. The pylon and local ablator will be redesigned to withstand the extreme heating conditions.

The following aircraft areas were of special interest following the flight:

a. Ball Nose: The preflight condition of the ablative material in the area of the ball nose is shown in figure 6. The top and bottom views of the ball nose area are shown in figures 7 and 8, respectively. The strips of additional heating noted in figure 7 appear to radiate aft from the ball nose b pressure ports as a result of flow-field interference. Flow-field interference effects were also noted on the lower surface, figure 8, aft of the reaction control rockets. The rocket motors had been plugged prior to the flight, however, the surface was not faired with ablative. These holes will remain plugged and sufficient ablative material will be applied to produce a smooth continuous flow past this area.

b. External Tank Door: The external tank door, shown in figure 9, remained extended in the airflow after the external tanks were ejected, resulting in additional ablation of material on and surrounding the door.

c. Antenna: A preflight view of the forward antenna is shown in figure 10. The antenna ablative material was repaired prior to flight 2-53. The postflight view of the antenna is shown in figure 11.

d. Protuberances: Typical effects of protuberances are shown in figure 12. These ranged from blisters and charring as a result of flow-field interference of the external tank cap to charring results of flow fields produced by extensions on the lower fuselage.

e. Windows: The right window was coated with ablative fogging about halfway back as shown in figure 13. The pilot reported visibility was good looking straight out, but was not acceptable looking forward. Some charring was noted on the canopy window leading edge as shown in figure 14.

f. Wing leading edges: View of the wing leading-edge ablative material after flight 2-53 are shown in figures 15 and 16. Char thickness for specified correlation points and the wing surfaces are shown in tables II, III, and IV.

g. Horizontal stabilizers: Char similar to that noted on the wing leading edges were observed on the horizontal stabilizers as shown in figure 17. Significant ablation was noted on the inboard fairing as shown in figure 17. The additional heating appears to have been a result of trapped flow created by the trough between the upper and lower surface of the horizontal stabilizer. Because of this localized heating, the ablative coating was charred on the back surface at the root rib and was lost due to rubbing on the side fairing. This trough will be filled with ablative material prior to the next ablative flight in order to eliminate this trough. Char thickness on the upper surface are tabulated in table V.

h. Upper Vertical: The ablative material on the upper vertical appears to have performed as planned. Ablation on the upper speed brake can be seen in figure 18. Char thickness measured on the upper vertical are shown in tables VI and VII.

i. Lower Vertical: The dummy ramjet shape flown under the lower vertical is shown in figure 19. The coating on the lower part of the ventral leading edge burned through approximately 140 seconds after launch at about maximum velocity. A very severe environment appears to have been caused by the interaction of the ventral-ramjet flow field. The resulting material erosion and high temperatures burned several holes in the ventral structure on and aft of the leading edges as shown in figures 20 through 23. An effort is currently being made to estimate heating rates and match calculated data with flight data for this area.

Ramjet Data were obtained from the dummy ramjet and pylon instrumentation from launch until about 25 seconds after shutdown. As a result of the shock impingement producing local aerodynamic heating and subsequent melting of pylon structure, acceleration and pressure data were lost as the severe heating burned through the ramjet instrumentation located in the pylon. The heating also produced some melting of the dummy ramjet skin and structure as shown in figure 24. Some aft pylon surface pressures and all speed brake surface pressures continued to function for the remainder of the flight.

The ramjet separated prematurely in the pattern during the turn to downwind as shown in figure 25. The following conditions existed at ramjet separation:

Altitude 32,000 feet

Velocity 1,000 feet per second

Mach number 1.02

Dynamic Pressure 420 psf

Angle of Attack 8 degrees

Normal Acceleration 1.6 "g"

Roll attitude -57 degrees

The parachute did not deploy due to heat damage and the dummy shape was severely damaged on impact as shown in figure 26. Analysis of the ramjet revealed three of the four explosive bolts retaining the ramjet had fired and the fourth had failed structurally. Indications are that the high-temperature environment caused the explosive bolts to ignite, however, the time at which this occurred is unknown. The ramjet instrumentation was burned and is not repairable. The flow-field cone probes were damaged beyond repair.

Preliminary data resembles that of previous flights, except for skin temperatures which are higher than previously obtained. A total failure of the high-response-side static was noted. Analysis of the data is continuing.

Fluid Total Temperature Probe

A fluidic total temperature probe and a shielded Pt - Pt Rh (10%) total temperature probe (as shown on figure 18) was flown on the leading edge of the upper vertical. The fluidic temperature probe and the Pt - Pt Rh (10%) both indicated total temperature in excess of 3000°R. A noise output signal from the fluidic probe, after engine shutdown, resulted in partial loss of data. The Pt - Pt Rh (10%) failed a few seconds after peak Mach number was reached. Refinement and analysis of the data are in progress.

Operational Discrepancies

During aircraft servicing, a persistent dribble of ammonia from the fuel jettison valve required that the fuel tank be pressurized to reseat the valve. Subsequent venting of the tank allowed the valve to back off again slightly, but the resulting dribble was quite small and considered acceptable. The valve will be replaced prior to the next flight.

A failure of the communications system within the pilot's helmet was noted after pilot entry. The helmet was replaced with a spare.

During the prelaunch tests of the stability augmentation system, the yaw channel failed to trip out on "Working Channel Test." Trouble had been experienced previously with the in-flight test unit. Since no inadvertent yaw channel tripouts had occurred and the yaw channel could be reset without indications of malfunction, it was decided to proceed. All channels appear to have functioned normally during the flight. The problem will be investigated when the aircraft becomes available for test.

Visual checks of the aircraft after touchdown indicate the ammonia external tank door remained open after the tank ejection. When the APU's were shut down, one or both of the external propellant transfer valves appeared to cycle. Photographic ground coverage indicated a slight vapor trail was tracking the X-15 after external tank ejection. Apparently one or both propellant transfer valves failed to close properly after the tanks were ejected, permitting the expulsion of propellants estimated as an insignificant amount, but sufficient to produce the light vapor trail noted on film. An investigation of this discrepancy is being conducted.

As a result of local aerodynamic heating at the ramjet pylon, three discrepancies were noted during or after the flight. The excessive heating propagated upward into the lower fuselage area, caused the over temperature indication of the engine peroxide compartment noted by the pilot at a velocity of 5000 feet per second, produced a failure of a helium control-gas line, which resulted in a condition from which a jettison of residual propellants could not be accomplished. The excessive heating eventually "cooked off" the explosive bolts attaching the dummy ramjet shape, permitting a premature separation. An investigation of the ramjet-pylon configuration is being conducted.

A preliminary review indicates the horizontal stabilizer control system may not have functioned properly during the high-temperature portion. The internal data are being investigated to determine whether a system malfunction has occurred.

Data Discrepancies

A preliminary review of data indicate the ball nose does not appear to have functioned properly during the high-temperature portion of the flight. Yaw pulses which yielded significant rates were not observed in angle-of-sideslip recordings during the high-temperature portions but were detected at lower temperatures. An investigation of this discrepancy is being conducted.

The Millikan camera installed in the 29-inch extension on the lower fuselage to view the lower ventral pylon-ramjet shape did not function because a malfunction of the thermal switch which activates the camera heater resulted in freezing of the camera.

The following internal data discrepancies were reported:

Parsons tape S/N 38: The ramjet side static data were not recorded due to an inoperative tape channel.

Parsons tape S/N 39: Trace 3 was noisy during portions of the flight.

0-25-36c: The thermocouple traces were too light for satisfactory telereader operation and the cal-step (Seq. #2) on channel 20 did not function.

0-19-36c: Timing lines and traces were too light. Ramjet spike data were not obtained on channels 2, 6, 7, and 8.

Longitudinal accelerometer (channel 15) was inoperative during the postflight.

The upper vertical stabilizer trace (channel 24) was too light to be read during portions of the flight and was off-scale during the postflight.

0-18-36c: Several traces were reported to be light.

The following channels were inoperative during postflight: 5, 7, 8, 10, 12, 21-25, 31, and 36.

0-26-36c: The following engine parameter channels were reported to be inoperative during the postflight: Pump speed, pump discharge, first and second stage, main chamber, and V port. The throttle micro switch recording does not appear to be functioning in a normal manner.

P-33-24: Timing lines in the center of the film are faint.

P-75-2: All channels were inoperative after time 44 seconds. Timing lines are readable only on edge of the film.

SV-5-3N: Pre- and postflight attitude zeros for corrections were not obtained.

P-1-4E: Airspeed and altitude traces became faint late in the flight.

SV-1-3N: Timing lines were too faint to be readable on the telereader. Inertial height traces jumps about 0.1-inch 90 seconds after launch.

(James R. Welsh, for)

E. J. Adkins, Chief

X-15 Research Planning Office

TABLE I

STABILITY AND CONTROL MANEUVERS

Maneuver Damper Angle of Attack
Degrees
Velocity Speed
Brakes

Yaw pulse Yaw off 7 6424 In

Yaw pulse Yaw low 5.5 5951 Out

Pitch pulse Pitch low 8 5670 Out

Push down - 10.5 - 2 5128-4990 Out

Pull up - 2.5 - 10 4957-4770 Out

Yaw pulse Yaw off 8 4700 Out

Yaw pulse Yaw off 6.5 4510 Out

Push down - 6.2 - 2 4450-4310 Out

Pull up - 2 - 10 4260-4140 Out

Yaw pulse Yaw off 4.5 3940 Out

Yaw pulse Yaw off 6 3490 Out

TABLE II

Post Flight Ablator Data: Flight 2-53-97

NASA Correlation Points

T/C No. Location Top Fwd Bot Aft Frt Lft Rt Char Pyrolysis

TF 2B Nose Panel F-2 .223 .222 .218 .229 .050 None Visible

04017 Nose Gear Door .171 .179 .182 .180 .086 None Visible

019 Canopy Ldg .532 .400 .344 .049 .187

Edge .513 .355 .419

TF18M Fwd Side Frg .133 .141 .137 .140 None None Visible

13050 Wing Ldg Edge .494 .422 .345 .175 .094

.533 .418 .367

13057 Wing Upr Surf .075 .076 .072 .080 None None Visible

13076 Wing Lwr Surf .213 .213 .213 .210 .042 None Visible

15017 Horiz Stab .622 .424 .369 .147 .095

Leading Edge .682 .423 .337

15018 H Stab Upr Sur .122 .142 .191 .193 .062 None Visible

15025 H Stab Lw Sur .165 .136 .158 .168 None None Visible

16020 Upr Vert Stab .473 .305 .322 .149 .129

Leading Edge .481 .302 .321

16025 U Ver Stab RH .287 .287 .276 .260 .005 None Visible

16018 Rt Up SB .193 .190 .204 .204 .075 None Visible

09001 Fuslg Rt side .111 .128 .094 .132 None None Visible

T20-22 Fuslg Lt side .172 .176 .176 .179 None None Visible

TABLE III