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The gust and draft velocities evaluated from records of NACA instruments installed in P-61C airplanes participating in thunderstorm flights at Clinton County Army Air Field, Ohio, from May 13 to May 29, 1947 are presented.
The time structure characteristics of agricultural airplane missions were studied by using records from NASA VGH flight recorders. Flight times varied from less than 3 minutes to more than 103 minutes. There was a significant reduction in turning time between spreading runs as pilot experience in the airplane type increased. Spreading runs accounted for only 25 to 29 percent of the flight time of an agricultural airplane. Lowering the longitudinal stick force appeared to reduce both the turning time between spreading runs and pilot fatigue at the end of a working day.
The effort required by the transport aircraft manufacturers to support the introduction of advanced composite materials into the fuselage structure of future commercial and military transport aircraft is investigated. Technology issues, potential benefits to military life cycle costs and commercial operating costs, and development plans are examined. The most urgent technology issues defined are impact dynamics, acoustic transmission, pressure containment and damage tolerance, post-buckling, cutouts, and joints and splices. A technology demonstration program is defined and a rough cost and schedule identified. The fabrication and test of a full-scale fuselage barrel section is presented. Commercial and military benefits are identified. Fuselage structure weight savings from use of advanced composites are 16.4 percent for the commercial and 21.8 percent for the military. For the all-composite airplanes the savings are 26 percent and 29 percent, respectively. Commercial/operating costs are reduced by 5 percent for the all-composite airplane and military life cycle costs by 10 percent.
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A semispan wing/body model with a powered highly loaded propeller has been tested to provide data on the propulsion installation drag of advanced propfan-powered aircraft. The model had a supercritical wing with a contoured over-the-wing nacelle. It was tested in the Ames Research Center's (ARC) 14-foot Transonic Wind Tunnel at a total pressure of 1 atm. The test was conducted at angles of attack from -0.5 to 4 deg at Mach numbers ranging from 0.6 to 0.8. The test objectives were to determine propeller performance, exhaust jet effects, propeller slipstream interference drag, and total powerplant installation drag. Test results indicated a total powerplant installation drag of 82 counts (0.0082) at a Mach number of 0.8 and a lift coefficient of 0.5, which is approximately 29 percent of a typical airplane cruise drag.
From 1977 to 1989, the NASA Langley Research Center conducted stall/spin flight tests of variations of four typical light airplanes. Each was equipped with an identical tail-mounted, spin-recovery parachute system. The system was used 29 times to arrest otherwise unrecoverable spins and was used at least twice on each airplane. The 10.5 ft diameter, ring-slot, spin-recovery parachute with 20 ft attachment lines and drag coefficient of .545 provided recovery from spins at angles of attack of 32 to 79 degrees and rotation rates of 122 to 261 degrees per second in less than 3 1/2 turns.
This bibliography, with abstracts, consists of 69 publications arranged in chronological order. The material may be useful to those interested in supersonic cruise fighter/penetrator/interceptor airplanes. Two pertinent conferences on military supercruise aircraft are considered as single items; one contains 37 papers and the other 29 papers. In addition, several related bibliographies are included which cover supersonic civil aircraft and military aircraft studies at the Langley Research Center. There is also an author index.
Fatal loss-of-control accidents have been directly related to in-flight airframe icing. The prototype system presented in this report directly addresses the need for real-time onboard envelope protection in icing conditions. The combination of prior information and real-time aerodynamic parameter estimations are shown to provide sufficient information for determining safe limits of the flight envelope during inflight icing encounters. The Icing Contamination Envelope Protection (ICEPro) system was designed and implemented to identify degradations in airplane performance and flying qualities resulting from ice contamination and provide safe flight-envelope cues to the pilot. The utility of the ICEPro system for mitigating a potentially hazardous icing condition was evaluated by 29 pilots using the NASA Ice Contamination Effects Flight Training Device. Results showed that real time assessment cues were effective in reducing the number of potentially hazardous upset events and in lessening exposure to loss of control following an incipient upset condition. Pilot workload with the added ICEPro displays was not measurably affected, but pilot opinion surveys showed that real time cueing greatly improved their awareness of a hazardous aircraft state. The performance of ICEPro system was further evaluated by various levels of sensor noise and atmospheric turbulence.
A flight test program was performed using the Boeing 757 flight research airplane to investigate the effect of noise from wing mounted engines on laminar boundary layer transition. An NLF glove was installed on the right wing panel just outboard of the engine. The extent of laminar flow on the glove was measured as a function of engine power setting for a range of flight conditions. A combination of surface and probe microphones was distributed over the upper and lower wing surfaces to measure sound spectra. The flight test program was completed in June 1985 and the results of preliminary analysis indicate that a maximum of about 29 percent of chord laminar flow was obtained on the upper surface and about 28 percent on the lower surface (at a high sideslip condition). The engine speed was varied from about 2600 (idle) to about 4500 (maximum continuous power) r/min. This produced changes in sound pressure level up to 20 dB on the lower surface. On the upper surface, the noise levels were independent of engine power but sensitive to airplane Mach number. No effect of engine power setting on upper surface transition location was observed, and only a small forward movement of the transition location on the lower surface was observed at the high power settings. Volume 1 of this report contains the program description and data analysis. Volume 2 is a compilation of all of the flight test data.
The papers in this compilation were presented at the NASA Symposium on "Supercritical Wing Technology: A Progress Report on Flight Evaluation" held at the NASA Flight Research Center, Edwards, Calif., on February 29, 1972. The purpose of the symposium was to present timely information on flight results obtained with the F-8 and T-2C supercritical wing configurations, discuss comparisons with wind-tunnel predictions, and project [ ] flight programs planned for the F-8 and F-III (TACT) airplanes.
An experimental investigation has been made in the Langley stability tunnel at low speed to determine the static longitudinal and lateral stability characteristics of a l/9-scale powered model of the Convair XFY-1 vertically rising airplane. Effects of thrust coefficient were investigated for the complete model and for certain components of the model. Effects of control deflections and of propeller-blade angle were investigated briefly for the complete model. Most of the tests were made through an angle-of-attack range from about -4 deg. to 29 deg, and the thrust-coefficient range was from 0 t o 0.7. In order to expedite distribution of these data to interested persons, no analysis of the data has been prepared for this report,
An analysis of 34 selected flight test data cases from a NASA flight program incorporating a natural laminar flow airfoil into partial wing gloves on the F-111 TACT airplane is given. This analysis determined the measured location of transition from laminar to turbulent flow. The report also contains the results of a boundary layer stability analysis of 25 of the selected cases in which the crossflow (C-F) and Tollmien-Schlichting (T-S) disturbance amplification factors are correlated with the measured transition location. The chord Reynolds numbers for these cases ranges from about 23 million to 29 million, the Mach numbers ranged from 0.80 to 0.85, and the glove leading-edge sweep angles ranged from 9 deg to 25 deg. Results indicate that the maximum extent of laminar flow varies from 56% chord to 9-deg sweep on the upper surface, and from 51% chord at 16-deg sweep to 6% chord at 25-deg sweep on the lower. The results of the boundary layer stability analysis indicate that when both C-F and T-S disturbances are amplified, an interaction takes place which reduces the maximum amplification factor of either type of disturbance that can be tolerated without causing transition.
An investigation has been conducted in the 27- by 27-inch preflight jet of the Langley Pilotless Aircraft Research Station at Wallops Island, Va., of the release characteristics of a dynamically scaled streamlined-type internally carried store from a simulated bomb bay at Mach numbers M(sub o) of 0.8, 1.4, and 1.98. A l/17-scale model of the Republic F-105 half-fuselage and bomb-bay configuration was used with a streamlined store shape of a fineness ratio of 6.00. Simulated altitudes were 3,400 feet at M(sub o) = 0.8, 3,400, and 29,000 feet at M(sub o) = 1.4, and 29,000 feet at M(sub o) = 1.98. At supersonic speeds, high pitching moments are induced on the store in the vicinity of the bomb bay at high dynamic pressures. Successful ejections could not be made with the original configuration at supersonic speeds at near sea-level conditions. The pitching moments caused by unsymmetrical pressures on the store in a disturbed flow field were overcome by replacing the high-aspect-ratio fin with a low-aspect-ratio fin that had a 30-percent area increase which was less subject to aeroelastic effects. Release characteristics of the store were improved by orienting the fins so that they were in a more uniform flow field at the point of store release. The store pitching moments were shown to be reduced by increasing the simulated altitude. Favorable ejections were made at subsonic speeds at near sea-level conditions.
As a continuation of an investigation of the release characteristics of an MB-1 rocket carried internally by the Convair F-106A airplane, six missile-bay baffle configurations and a rocket end plate have been investigated in the 27- by 27-inch preflight jet of the NASA Wallops Station. The MB-1 rocket used had retractable fins and was ejected from a missile bay modified by the addition of six different baffle configurations. For some tests a rocket end plate was added to the model. Dynamically scaled models (0.04956 scale) were tested at a simulated altitude of 22,450 feet and Mach numbers of 0.86, 1.59, and 1.98, and at a simulated altitude of 29,450 feet and a Mach number of 1.98. The results of this investigation indicate that the missile-bay baffle configurations and the rocket end plate may be used to reduce the positive pitch amplitude of the MB-1 rocket after release. The initial negative pitching velocity applied to the MB-1 rocket might then be reduced in order to maintain a near-level-flight attitude after release. As the fuselage angle of attack is increased, the negative pitch amplitude of the rocket is decreased.
The F404-GE-400 engine powered F/A-18A High Alpha Research Vehicle (HARV) was used to examine the characteristics of inlet airflow during rapid aircraft maneuvers. A study of the degree of similarity between inlet data obtained during rapid aircraft maneuvers and inlet data obtained at steady aerodynamic attitudes was conducted at the maximum engine airflow of approximately 145 Ibm/sec using a computer model that was generated from inlet data obtained during steady aerodynamic maneuvers. Results show that rapid-maneuver inlet recoveries agreed very well with the recoveries obtained at equivalent stabilized angle-of-attack conditions. The peak dynamic circumferential distortion values obtained during rapid maneuvers agreed within 0.01 units of distortion over the 10 - 38 degree angle of attack range with the values obtained during steady aerodynamic maneuvers while similar agreement was found for the peak dynamic radial distortion values up to 29 degrees angle-of-attack. Exceedences of the rapid-maneuver peak dynamic circumferential distortion values relative to the peak distortion model values at steady attitudes occurred only at low or negative angles of attack and were inconsequential from an engine-stability assessment point of view. The results of this study validate the current industry practice of testing at steady aerodynamic conditions to characterize inlet recovery and peak dynamic distortion levels.
A leading edge flap (16) for supersonic transport airplanes is disclosed. In its stowed position, the leading edge flap forms the lower surface of the wing leading edge up to the horizontal center of the leading edge radius. For low speed operation, the vortex leading edge flap moves forward and rotates down. The upward curve of the flap leading edge triggers flow separation on the flap and rotational flow on the upper surface of the flap (vortex). The rounded shape of the upper fixed leading edge provides the conditions for a controlled reattachment of the flow on the upper wing surface and therefore a stable vortex. The vortex generates lift and a nose-up pitching moment. This improves maximum lift at low speed, reduces attitude for a given lift coefficient and improves lift to drag ratio. The mechanism (27) to move the vortex flap consists of two spanwise supports (24) with two diverging straight tracks (64 and 68) each and a screw drive mechanism (62) in the center of the flap panel (29). The flap motion is essentially normal to the airloads and therefore requires only low actuation forces.
The Propulsion Flight Test Fixture at the NASA Dryden Flight Research Center is a unique test platform available for use on the NASA F-15B airplane, tail number 836, as a modular host for a variety of aerodynamics and propulsion research. The first experiment that is to be flown on the test fixture is the Channeled Centerbody Inlet Experiment. The objectives of this project at Dryden are twofold: 1) flight evaluation of an innovative new approach to variable geometry for high-speed inlets, and 2) flight validation of channeled inlet performance prediction by complex computational fluid dynamics codes. The inlet itself is a fixed-geometry version of a mixed-compression, variable-geometry, supersonic in- let developed by TechLand Research, Inc. (North Olmsted, Ohio) to improve the efficiency of supersonic flight at off-nominal conditions. The concept utilizes variable channels in the centerbody section to vary the mass flow of the inlet, enabling efficient operation at a range of flight conditions. This study is particularly concerned with the starting characteristics of the inlet. Computational fluid dynamics studies were shown to align well with analytical predictions, showing the inlet to remain unstarted as designed at the primary test point of Mach 1.5 at an equivalent pressure altitude of 29,500 ft local conditions. Mass-flow-related concerns such as the inlet start problem, as well as inlet efficiency in terms of total pressure loss, are assessed using the flight test geometry.
In June 2019, a full-scale crash test of a Fokker F28 Fellowship aircraft will be conducted at NASA Langley Research Center?s Landing and Impact Research (LandIR) Facility. The F28 is a high-performance twin-turbo fan narrow-body aircraft with seating in a 3+2 configuration. The MK4000 variant, used in this test, is capable of carrying up to 85 passengers on medium range routes. The F28 was first type certified by the Federal Aviation Administration (FAA) in 1969 and now the majority of the F28 fleet has retired from service in the United States. In 2016, the FAA and NASA Langley Research Center (LaRC) signed an interagency agreement for conducting a research program to obtain test data that will support the development of airframe level crash requirements for transport category airplanes [1]. The objectives of the full-scale crash test can be divided into six categories: (1) To compare and contrast responses in identical aircraft undergoing vertical only to combined vertical and horizontal loading conditions, (2) To examine the effects of horizontal loading on aircraft structure during a crash event, (3) To generate data for the use in calibration of computer modelling efforts, (4) To generate data from onboard Anthropomorphic Test Devices (ATDs) for the evaluation of injury, (5) To obtain data from experimental seats, and (6) To obtain data from new and novel ATD designs including Warrior Injury Assessment MANikin (WIAMan) [2], Test device for Human Occupant Response (THOR), and other newly developed child ATDs. The focus of this presentation will be to document finite element model development of the full-scale F28 aircraft and to present preliminary test-analysis predictions. NASA obtained the full-scale F28 aircraft, plus three fuselage sections and two sets of wings with funding through the NASA Aviation Safety Program in 1998. In addition to the hardware, NASA purchased a full NASTRAN model of the airframe that had been developed by the Dutch manufacturer, Fokker. Beginning in 2016, the NASTRAN model was converted to LS-DYNA? [4, 5] format and modified by combining parts, adding missing parts of internal structure, and including ballast for loading weights, etc. It contains: 89,223 nodes; 24,065 beam elements; 55,404 shell elements; 29,044 solid elements; 740 parts; 80 material definitions; and, 46 Constrained Nodal Rigid Bodies (CNRBs). The aircraft model will be executed in LSDYNA to simulate the test article impact onto a 2-ft.-high bed of soil under combined velocity conditions of 70-ft/s forward and 30-ft/s vertical velocity. Pre-test simulation predictions will be generated and correlated with test data.