- SuperSonic Natural Laminar Flow (SS-NLF)
- Gust Monitoring and Aeroelasticity Experiment
- X-33 Thermal Protection Systems (TPS) Rain Erosion Studies
- X-33 and Space Shuttle Durability Studies
- Turbulent Skin Friction Reduction Experiment
- Shock Wave Location Sensor
- Aerostructures Test Wing
- F-5 Shaped Sonic Boom Demonstration
SuperSonic Natural Laminar Flow (SS-NLF)
The primary goal of the SuperSonic Natural Laminar Flow experiment was to investigate the extent and stability of natural laminar flow at supersonic speeds on a specially designed airfoil. Additional goals were: a) to refine the technique of using infra-red (IR) thermography to identify and locate transition fronts and shocks, and obtain flight data to validate analytical techniques for aerodynamic and thermal predictions.
Using an aircraft mounted infra-red camera, laminar flow was measured on a test wing mounted on the centerline pylon of an F-15B aircraft. An infra-red camera was used to measure surface temperatures that change with the different boundary layer states. The surface beneath the turbulent boundary layer was warmer due to the higher convection of the turbulent layer. The laminar flow, for these conditions, was darker and the turbulent flow lighter, as shown in the infra-red image to the right.
|Supersonic Natural Laminar Flow test article mounted on F-15B aircraft.||Infrared image of Supersonic Natural Laminar Flow test fixture.|
The test wing was fabricated from aluminum with an insulating layer covering all but the first 1 to 2 inches of the leading and trailing edges. A splitter plate was formed over the test wing to minimize disturbances from the bottom of the aircraft effecting the test surface. Laminar flow was obtained to approximately 80% of the inner 2/3. The laminar flow was able to penetrate weak shock waves, but was typically terminated by strong shock waves. The strongest shock wave appeared to emanate from the camera pod located outboard on the armament rail.
SuperSonic Natural Laminar Flow was a joint effort between NASA, Directed Technologies Inc., and Reno Aeronautical Corporation.
Gust Monitoring and Aeroelasticity Experiment
The Gust Monitoring and Aeroelasticity (GMA) experiment was initiated by University of California, Los Angeles (UCLA) graduate students under the UCLA/NASA Center for Flight Systems Research. The primary purpose of the Gust Monitoring and Aeroelasticity flight experiment was to measure the aeroelastic response of a small test wing, and to monitor in-flight gust intensity using a customized laser-based device.
The small test wing and a low-power laser were flown attached to the aft/left panel of the Flight Test Fixture-II test article. The small, flexible 18-inch long, 4-inch chord, test wing was instrumented with accelerometers and strain gages to measure the dynamic response of the wing and Flight Test Fixture-II. Video cameras mounted on the underside of the aircraft recorded the movement of the test wing. The Gust Monitoring and Aeroelasticity experiment was flown on the F-15B / Flight Test Fixture-II at speeds of up to Mach 0.80 and altitudes up to 20,000 feet over a series of 4 flights.
Flight data was used to investigate the relationship of the dynamic response of a flexible wing to in-flight maneuvers and atmospheric wind gusts. In addition, the experiment tested a new gust-monitoring device based on measurements of the forward scattering of a laser beam. New aeroelastic modeling techniques are being developed and will be tested using data collected during this experiment. These modeling techniques are being developed to aid in the synthesis of active gust alleviation and flutter suppression control algorithms.
X-33 Thermal Protection Systems (TPS) Rain Erosion Studies
Using the F-15B / Flight Test Fixture-II as a testbed, Dryden completed testing of the X-33 Reusable Launch Vehicle (RLV) thermal protection systems materials for durability at flight velocities through adverse weather, much like the tests conducted on Dryden's F-104 Flight Test Fixture for space shuttle tiles many years ago. A primary threat to thermal protection systems materials during ascent and landing is impact with rain drops, cloud droplets or ice crystals. The proposed requirement of "launch on demand" for the X-33 required that thermal protection systems materials resist damages from these high velocity impacts.
In September of 1985, using the F-104 Flight Test Fixture as a testbed, Dryden began testing space shuttle thermal protection systems tiles using a similar test fixture, incorporating both new tiles and tiles that had flown in space on Columbia. Today's generation of advanced ceramic and metallic thermal protection systems materials designed by NASA and Rockwell International are more durable. These advanced thermal protection systems materials were installed on a nose section, built by Rockwell International, that was adapted to the forward section of the Flight Test Fixture-II. thermal protection systems tiles were mounted on the nose section at several angles to simulate impacts to the thermal protection systems materials on the X-33 Reusable Launch Vehicle. These thermal protection systems materials were flown through adverse weather conditions over a range of flight velocities.
A particle measurement probe, the same used in prior F-104 tests, was installed under the Flight Test Fixture-II to measure drop size, distribution and concentration. Edwards weather radar data was also used to predict rain location and quantity during flight testing. Video cameras on the F-15B were used to document damage to the thermal protection systems materials. Dryden conducted these tests with Rockwell International in support of the X-33 Reusable Launch Vehicle and Marshall Space Flight Center.
X-33 and Space Shuttle Durability Studies
The Flight Test Fixture-II has flown many flight experiments over the past several years and can be modified to support a variety of research requirements. For example, the X-33 program requested assistance in exposing test articles to shear and impinging shock loads in a flight environment as part of the flight qualification of the X-33 thermal protection systems materials. X-33 thermal protection systems materials for this experiment ranged from metallic panel materials supplied by BF Goodrich to a variety of advanced flexible reusable surface insulation (AFRSI) specimens supplied by NASA Ames Research Center. Transition seals and flight test instrumentation islands were also incorporated into the test articles to demonstrate the durability of these components. Some of the test articles were thermally cycled in an arc-jet tunnel prior to flight test in order to simulate thermal loads expected on the X-33 vehicle.
The two forward left side panels on the Flight Test Fixture-II were replaced by a large carrier plate in order to simplify the installation of the various thermal protection systems test articles. This concept was developed to allow for quick configuration changes between research flights. Test articles were installed in the various quadrants of the carrier plate depending on the desired configuration for each flight. Forward test articles were generally used to look at the effect of shock impingement loads previously identified at this location at transonic speeds. Test articles in the aft locations were used to document the effect of shear loads.
Six configurations were flight tested at a maximum Mach number of 1.4 and dynamic pressures as high as 790 lbs/sq.ft. Flight test were conducted as low as 5,000 feet to obtain the higher shear loads and as high as 35,000 feet for supersonic testing. Surface pressures were obtained to document flow conditions and loads on the test articles. In addition, in-flight video and detailed pre- and post-flight photos were used to document the condition of all test articles. This highly successful flight test series was completed in May 1998.
Several months later, the Shuttle External Tank Project from Marshall Space Flight Center saw the results of the X-33 test and requested use of the same carrier plate to expose test articles of the shuttle's external tank insulation to a simulated shuttle launch environment up to Mach 1.5 and 60,000 ft. Six test articles were flown to simulate the thrust panel rib structure and foam where the shuttle solid rocket boosters are attached to external tank. Several types of foam insulation configurations and rib orientations were tested using similar instrumentation and video documentation. This flight test series was successfully completed in January 1999 in less than two weeks.
The X-33 and Shuttle External Tank Projects were both able to gain significant benefit from the flight test results obtained using the F-15B Flight Test Fixture-II with a carrier plate as a research platform. Flight results were obtained quickly and efficiently, and provided valuable data toward flight qualification with an increased understanding of the durability of these materials in an actual flight environment.
Turbulent Skin Friction Reduction Experiment
Results from preliminary research conducted by Eidetics Corp. and others indicated that turbulent skin friction could be reduced by heating the surface with electric resistance heaters. These results, however, were obtained at very low Reynolds and Mach numbers. The F-15B/Flight Test Fixture-II provided a means to determine how much skin friction could be reduced at typical transport aircraft flight conditions and the power required. The second part of this Small Business Innovative Research (SBIR) Phase II by Eidetics Corp. extended the experiment to the forward fuselage of a T-39 Sabreliner. Conducting the experiment on the F-15B/Flight Test Fixture-II first provided researchers with confidence in the adhesive and installation technique prior to the resistance heaters being installed on the Sabreliner.
The F-15B/Flight Test Fixture-II experiment required the installation of seven 10" x 10" resistance heaters, a total pressure boundary layer rake, a total temperature boundary layer rake and various pressure and temperature and other sensors. In all, 156 parameters were recorded on each flight. Two rake configurations and fourteen heater combinations were tested over three altitudes and three Mach numbers. Three flights were required to test all of these combinations. Skin friction over the panels was determined using a momentum deficit technique.
Shock Wave Location Sensor
The F-15B Flight Test Fixture-II was used in the Fall of 1996 to test a newly developed aerodynamic shock wave location sensor. The flush surface mounted sensor was developed by Tao Systems. The device utilizes a multi-element hot-film sensor array, a multi-channel constant voltage anemometry system, and specially designed flow diagnostic software to identify the location of the shock wave and its dynamic characteristics.
Flight data were obtained with the sensor installed on a modified NACA 0021 airfoil that was attached to the left side of the Flight Test Fixture-II. The sensor successfully identified the transonic shock phenomena using both minimum voltage output and phase reversal analyses. This experiment was a significant step in developing a real-time shock location sensor. Wind tunnel studies are currently in progress to better understand the highly-complex off-surface flow phenomena which influence the measurements of the surface mounted hot-film sensors. Future flight tests on the F-15B Flight Test Fixture-II are anticipated to demonstrate the real-time capability of the system.
Aerostructures Test Wing
The Aerostructures Test Wing (ATW) flight experiment conducted by Dryden during April and May of 2001 successfully demonstrated a new software data analysis tool, the flutterometer, which is designed to increase the efficiency of flight flutter testing.
The experiment consisted of an 18-inch carbon fiber test wing with surface-mounted piezoelectric strain actuators. The test wing was mounted on a special ventral flight test fixture and flown on Dryden's F-15B Research Testbed aircraft.
Five flights consisted of increasing speeds and altitudes leading to the final test point of Mach .85 at an altitude of 10,000 feet. At each Mach and altitude, stability estimations of the wing were made using accelerometer measurements in response to the piezoelectric actuator excitation. The test wing was intentionally flown to the point of structural failure, resulting in about a third of the 18-inch wing breaking off. This allowed engineers to record the effectiveness of the flutterometer over the entire regime of flutter testing, up to and including structural failure.
The actuators were moved at different magnitudes and frequency levels to induce wing vibrations and excite the dynamics during flight. The Aerostructures Test Wing experiment represents the first time that piezoelectric actuators were used during a flight flutter test.
Potential benefits of this research include reduced time and cost associated with aircraft certification by lowering the number of flights required to clear a new or modified aircraft for flight, and provision of a structural dynamics database for industry and university flutter research.
The flutterometer is an on-line software tool that was loaded on computers in Dryden's control room. Using the flutterometer, flight data can be analyzed immediately to determine the stability properties of aircraft in flight, helping predict the flight conditions at which the onset of flutter may occur. In this way, the flight envelope of an aircraft can be found more quickly and safely than using traditional approaches. The Aerostructures Test Wing experiment was the first time the flutterometer was used on a flight system that actually experienced flutter.
Dryden engineers developed the flight data algorithms that made the flutterometer concept a reality. NASA was awarded a patent for the flutterometer. Its software program combines the strengths of analytical predictions and on-line estimation methods in the development of a flutterometer concept. The flutterometer software has previously been evaluated using simulations and wind tunnels along with flight data from several aircraft types including NASA Dryden's F-18 Systems Research Aircraft.
Flutter is the rapid and self-excited vibration of wings, tail surfaces, and other aircraft parts that can damage or destroy an aircraft component. Flutter is caused by the flow of air across the surface of the structure. Effectively, the aerodynamic forces couple with the structural bending and twisting to result in the destructive vibration. Flight flutter testing is the process of determining a flight envelope within which an aircraft is safe to operate. Traditional approaches for flight flutter testing do not accurately predict the onset of instability so this testing is a very time-consuming and expensive process.
F-5 Shaped Sonic Boom Demonstration
NASA is part of a team seeking quieter sonic booms. NASA's F-15B Research Testbed aircraft recently flew in the supersonic shock wave of a U.S. Navy F-5E in support of the F-5 Shaped Sonic Boom Demonstration (SSBD) project, part of the Defense Advanced Research Projects Agency's (DARPA) Quiet Supersonic Platform (QSP) program.
The flights originated from the NASA Dryden Flight Research Center at Edwards, Calif.
Four flights were flown in order to measure the F-5E's near-field (close-up) sonic boom signature at Mach 1.4, during which more than 50 shockwave patterns were measured at distances as close as 100 feet below the F-5E.
The F-15B's specially-instrumented noseboom recorded static pressure measurements while flying behind and below the F-5E. This provided a baseline measurement of the F-5E's sonic boom characteristics. Differential Global Positioning System (GPS) receivers on both aircraft yielded relative aircraft position.
Northrop Grumman Corporation's Integrated Systems sector of El Segundo, Calif., intends to modify the F-5E aircraft into a Shaped Sonic Boom Demonstrator in an effort to reduce its sonic boom. The U.S. Navy aircraft is based at Fallon Naval Air Station in Fallon, Nevada. The data provided by Dryden's F-15B is assisting Northrop Grumman's and DARPA's efforts, as well as helping assess SSBD flight test techniques.
In addition to the airborne data collected by the F-15B, sonic boom data was gathered on the ground by two Dryden-developed Boom Amplitude and Direction Sensors (BADS) in order to obtain ground-level sonic boom signature data. Twenty-five sonic booms from the F-5E and F-15B were recorded.
Dryden has expertise in air and ground-based sonic boom measurement techniques, having accomplished several sonic boom studies over the years. In 1995, Dryden's F-16XL-1 aircraft probed the shockwave of one of Dryden's SR-71 aircraft.
The flight data show fine details unseen in the preflight predictions. Based on these details, the Computational Fluid Dynamic (CFD) grid density was increased. Preliminary flight data agree well with the CFD predictions over most of the region, with an adjustment needed to the predictions in the region of the engine inlet. These flight data allow the QSP team to validate prediction tools to design aircraft with lower sonic booms.
DARPA and Northrop Grumman plan to fly the F-5E in the fall with a special fairing designed to reduce the aircraft's sonic boom. Dryden's F-15B will again fly in the Shaped Sonic Boom Demonstrator's shock waves to record changes produced by the F-5E modifications.