The current research examines the possibility of using recirculation filters from aircraft to document the nature of air-quality incidents on aircraft. These filters are highly effective at collecting solid and liquid particulates. Identification of engine oil contaminants arriving through the bleed air system on the filter was chosen as the initial focus. A two-step study was undertaken. First, a compressor/bleed air simulator was developed to simulate an engine oil leak, and samples were analyzed with gas chromatograph-mass spectrometry. These samples provided a concrete link between tricresyl phosphates and a homologous series of synthetic pentaerythritol esters from oil and contaminants found on the sample paper. The second step was to test 184 used aircraft filters with the same gas chromatograph-mass spectrometry system; of that total, 107 were standard filters, and 77 were nonstandard. Four of the standard filters had both markers for oil, with the homologous series synthetic pentaerythritol esters being the less common marker. It was also found that 90% of the filters had some detectable level of tricresyl phosphates. Of the 77 nonstandard filters, 30 had both markers for oil, a significantly higher percent than the standard filters.
An analytical investigation to model the manner in which pilots perceive and utilize visual, proprioceptive, and vestibular cues in a ground-based flight simulator was undertaken. Data from a NASA Ames Research Center vertical motion simulator study of a simple, single-degree-of-freedom rotorcraft bob-up/down maneuver were employed in the investigation. The study was part of a larger research effort that has the creation of a methodology for determining flight simulator fidelity requirements as its ultimate goal. The study utilized a closed-loop feedback structure of the pilot/simulator system that included the pilot, the cockpit inceptor, the dynamics of the simulated vehicle, and the motion system. With the exception of time delays that accrued in visual scene production in the simulator, visual scene effects were not included in this study. Pilot/vehicle analysis and fuzzy-inference identification were employed to study the changes in fidelity that occurred as the characteristics of the motion system were varied over five configurations. The data from three of the five pilots who participated in the experimental study were analyzed in the fuzzy-inference identification. Results indicate that both the analytical pilot/vehicle analysis and the fuzzy-inference identification can be used to identify changes in simulator fidelity for the task examined.
An experiment examined how visual scene and platform motion variations affected a pilot's ability to perform altitude changes. Pilots controlled a helicopter model in the vertical axis and moved between two points 32-ft apart in a specified time. Four factors were varied: visual-scene spatial frequency, visual-scene background, motion-filter gain, and motion-filter natural frequency. Drawing alternating black and white stripes of varying widths between the two extreme altitude points varied visual-scene spatial frequency. The visual-scene background varied by either drawing the stripes to fill the entire field of view or by placing the stripes on a narrow pole with a natural sky and ground plane behind the pole. Both the motion-filter gain and natural frequency were varied in the motion platform command software. Five pilots evaluated all combinations of the visual and motion variations. The results showed that only the motion-filter natural frequency and visual-scene background affected pilot performance and their subjective ratings. No significant effects of spatial frequency or motion system gain were found for the values examined in this tracking task. A previous motion fidelity criterion was found to still be a reasonable predictor of motion fidelity.
In airborne applications the high degree of reliability is required. Typical methods for assuring the reliability are development guides (DO178), testing, quality checking (AS9100) and certification processes. Despite these processes provide high level of safety and reliability of the products, the electronic devices can fail from different reasons. Also, one of the main present problems today is incompatibility of communication interfaces of so called smart sensors. This article proposes fusion of smart sensor standard IEEE 1451 and information necessary for sensor self-validation of sensor output. In this paper estimation of the feedback magnitude measured by an intelligent servomechanism system is presented. This system is being developed for a specific UAV application. The necessary data are saved within the TEDS memory of the extended IEEE 1451 standard. In the paper, following topics are discussed: methods of magnitudes validation and estimation, standards of smart sensors, the servopsilas implementation and employment in an UAV project, algorithms and data stored in the TEDS memory.
Several conceptual hypersonic research airplanes, designed within the constraints of a B-52 launch aircraft, have been studied experimentally and analytically at Mach numbers from 0.2 to 6.0. Vehicles built to these criteria for Mach 6 cruise were shown to be feasible, if careful attention was paid to the low speed lift, drag, and high angle of attack stability to assure successful landings and transonic pitch angle maneuvers. The integrated scramjet engine drag was high at subsonic speeds and appears to be constant with Reynolds number. The variable geometry airfoil used previously to improve directional stability was shown to be equally adaptable to the improvement of longitudinal stability. The vortex lattice theory gave good subsonic predictions of lift, drag due to lift, and pitching moments. Wind tunnel tests must be relied on for the drag at zero lift, trim, static margins and lateral-directional stability. The Gentry Hypersonic Arbitrary Body Program gave good predictions of the trends of lift, drag, and pitching moments with angle of attack at Mach numbers above 3, but the level of the values were not consistently predicted. No currently available theory or program gave accurate predictions of directional stability or dihedral effects at hypersonic speeds.
An experimental program was conducted with an NACA 0012 airfoil to investigate the effect of Reynolds number on the aerodynamic coefficients with and without a leading-edge simulated ice shape. Six Reynolds numbers from 0. 36 to 3. 36 multiplied by 10**6 were investigated at both positive and negative angles of attack through stall. Values of DELTA C//d, DELTA C// L-SCRIPT , DELTA C// L-SCRIPT //m//a//x, and DELTA C(m//c/////4) were measured for the no-ice/generic-ice airfoil configurations. The experimental values have shown that the addition of the generic ice shape causes premature stall with a considerable reduction in C(// L-SCRIPT //m//a//x) and stall angle of attack; an increase in the drag values of 120-200% compared to the clean airfoil values; and a significant increase in C(m//c/////4) resulting in positive values of dC//m/dC// L-SCRIPT . However, the aerodynamic coefficients of the airfoil with the leading-edge simulated ice showed little dependence on Reynolds number throughout the range tested, unlike that of the clean airfoil configuration.
The effect of a simulated glaze-ice accretion on the aerodynamic performance of a NACA 0012 airfoil was studied experimentally. Two ice shapes were tested: one from an experimentally measured accretion, and one from an accretion predicted using a computer model given the same icing conditions. Lift, drag, and pitching moment were measured for the airfoil with both smooth and rough ice shapes. The ice shapes caused large lift and drag penalties, primarily due to large separation bubbles. Surface pressure distributions clearly showed the regions of separated flow. The aerodynamic performance of the two shapes compared well at positive, but not negative, angles of attack.
Local convective heat transfer coefficients were measured from a smooth NACA 0012 airfoil having a chord length of 0.533 m. Flight data were taken for the smooth airfoil at Reynolds numbers based on chord in the range 1.24 to 2.50 million and at various angles of attack up to 4 deg. During these flight tests, the freestream velocity turbulence intensity was found to be very low. Wind tunnel data were acquired in the Reynolds number range 1.20 to 4.52 million and at angles of attack from -4 to +8 deg. The turbulence intensity in the IRT was 0.5-0.7 percent with the cloud-generating sprays off. A direct comparison between the results obtained in flight and in the IRT showed that the higher level of turbulence intensity in the IRT had little effect on the heat transfer for the lower Reynolds numbers but caused a moderate increase in heat transfer at the higher Reynolds numbers. Turning on the cloud-generating spray nozzle atomizing air in the IRT did not alter the heat transfer. The present data were compared with leading-edge cylinder and flat plate heat transfer correlations that are often used to estimate airfoil heat transfer in computer codes.
Local heat transfer measurements from a roughened NACA 0012 airfoil were successfully obtained in flight and in the NASA Lewis icing research tunnel using the method and apparatus described in this work. Major conclusions resulting from this study are as follows. The addition of roughness to the airfoil surface drastically increased the heat transfer downstream of stagnation. The roughness elements disturbed the laminar boundary-layer flow and in some cases caused a transition to turbulent flow. Compartison of the flight and tunnel roughened surface data showed that the general effect of increased turbulence was a slight increase in heat transfer, especially at the higher Reynolds numbers. Generally, the roughened surface airfoil cases showed the suction side heat transfer monotonically increasing with angle of attack.
Experimental oil-flow and tuft patterns and vapor-screen flow-visualization data were obtained on a cambered wing model at Mach = 1.62 for an angle-of attack range of 0-14 deg. These data were used as flow diagnostic tools along with surface-pressure and force data and full-potential theory calculations. A large separation bubble was found on the lower wing surface at low angle of attack. The high-angle-of-attack flowfield was characterized by a large attached-flow leading-edge expansion followed by a crossflow shock. At alpha = 14 deg the crossflow shock apparently induced discrete regions of streamwise separated flow, which were clearly indicated in the vapor-screen and oil-flow photographs.
Core noise from a YF-102 high bypass ratio turbofan engine was investigated through the use of simultaneous measurements of internal fluctuating pressures and far field noise. Acoustic waveguide probes, located in the engine at the compressor exit, in the combustor, at the turbine exit, and in the core nozzle, were employed to measure internal fluctuating pressures. Spectra showed that the internal signals were free of tones, except at high frequency where machinery noise was present. Data obtained over a wide range of engine conditions suggest that below 60% of maximum fan speed the low frequency core noise contributes significantly to the far field noise.
Cloud-to-ground (CG) lightning strike data on the NASA F-106B research aircraft obtained during the 1984-86 storm seasons in the vicinity of Wallops Island, Virginia, are analyzed. The results suggest that CG strikes may represent a significant portion of the total number of lightning strikes encountered by the aircraft at altitudes below 6 km. It is unlikely that an aircraft encounters the first return stroke of the CG flash. The current values of the CG strikes are not different from currents in other types of strikes.
A flight research program was conducted at NASA Langley Research Center to apply vapor-screen technology, used in wind tunnels, to aircraft. The purpose was to obtain qualitative and quantitative information about near-field vortex flows above the wings of fighter aircraft and ascertain the effects of Reynolds and Mach number over an angle-of-attack range. In order to use the vapor-screen technology, a complete set of systems had to be developed and flight qualified on an available aircraft that satisfied vortex-flow generation requirements. The F-106B was such an aircraft. To help establish the Reynolds and Mach number effects on vortex systems, two types of maneuvers were conducted. The first was at 1-g, consisting of various constant altitude level flights at angles of attack up to 23 deg and Mach number near 0.4 The second was at a high-g value near 5, consisting of a spiral descent at high angle of attack and Mach number near 0.8. Both kinds of maneuvers were done at night to ensure sufficient contrast between the ambient light and the illuminated vortex system. This paper describes the systems, assesses their performance, and summarizes the scientific results obtained from the photographic records.
This paper discusses and evaluates the test measurement techniques used to determine the laminar-to-turbulent boundary-layer transition location in the F-14 variable-sweep transition flight experiment (VSTFE). The main objective of the VSTFE was to determine the effects of wing sweep on the laminar-to-turbulent transition location at conditions representative of transport aircraft. Four methods were used to determine the transition location: (1) a hot-film anemometer system, (2) two boundary-layer rakes, (3) surface pitot tubes, and (4) liquid crystals for flow visualization. Of the four methods, the hot-film anemometer system was the most reliable indicator of transition.
A description of the basic design concept, hardware design, and flight evaluation of a Variable Stability System (VSS) installed on the NASA ARC X-14B is presented. The NASA ARC X-14B is a twin-engine, single-seated VTOL aircraft. The VSS is unique in that it employs a general purpose airborne digital computer as an integral part of the hybrid model following flight control system. The system design, analysis and testing phases are discussed in the paper from the application of optimal control techniques in the preliminary design of the system, through the flight demonstration of the VSS hardware.
An analysis is presented of high resolution wind profile measurements recorded at the NASA 150-m ground winds tower facility, showing wind speed shear frequency and magnitude distributions for six vertical layers of the atmosphere and one vertical distance. Vertical wind shear is defined as the change of wind speed with height, and its magnitudes were derived by algebraically subtracting lower level wind speeds from those of higher levels and dividing the distance between levels. Horizontal wind shear is understood to be change of wind speed with horizontal distance, and its magnitudes were derived by algebraically subtracting the wind speed at a short tower from that at a tall one and dividing by the distance between towers.
The transonic flowfield around an F-16A fighter configuration at a moderate incidence angle is simulated by solving the Navier-Stokes equations on a single-block grid. The numerical solution matches experimental freestream conditions with a Mach number of 0.85, 16-deg angle of attack, and a characteristic Reynolds number of 12.75 million. MacCormack's explicit algorithm is used in conjunction with a local time step and consecutive mesh refinement procedure to accelerate numerical convergence. The Baldwin-Lomax algebraic model provides turbulent closure. Computed surface pressure distributions and the aircraft lift coefficient compare favorably with wind tunnel data. The drag coefficient in the simulation overpredicts the experimental value by 8%. The analysis of results details the interactions of the strake vortex and the wing root vortex in an upwash flow. This discussion also focuses on the importance of grid topology and surface definition in the computation of flows at high angle of attack.
The TranAir computer code solves the full-potential equation for transonic flow by combining a rectangular box of flowfield grid points with networks of surface panels. Complex geometries are easily represented since surface-conforming flowfield grids are not used. Wing pressures predicted by TranAir are compared to wind-tunnel data at alpha equals 4 deg for M// infinity equals 0. 6 and 0. 9. The higher Mach number condition produces supercritical flow and demonstrates TranAir's shock-capturing capability.
A series of ground-based studies have been conducted to develop actuated forebody strake controls for flight test evaluations using the NASA F-18 High-Alpha Research Vehicle. The actuated forebody strake concept has been designed to provide increased levels of yaw control at high angles of attack where conventional rudders become ineffective. Results are presented from tests conducted with the flight-test strake design, including static and dynamic wind-tunnel tests, transonic wind-tunnel tests, full-scale wind-tunnel tests, pressure surveys, and flow visualization tests. Results from these studies show that a pair of conformal actuated forebody strakes applied to the F-18 HARV can provide a powerful and precise yaw control device at high angles of attack. The preparations for flight testing are described, including the fabrication of flight hardware and the development of aircraft flight control laws. The primary objectives of the flight tests are to provide flight validation of the groundbased studies and to evaluate the use of this type of control to enhance fighter aircraft maneuverability.
As part of the NASA High Alpha Technology Program, fine-grid Navier-Stokes solutions have been obtained for flow over the fuselage forebody and wing leading-edge extension of the F/A-18 high alpha research vehicle at large incidence. The resulting flows are complex, and exhibit crossflow separation from the sides of the forebody and from the leading edge extension. A well-defined vortex pattern is observed in the leeward-side flow. Results obtained for laminar flow good agreement with flow visualizations obtained in ground-based experiments. Further, turbulent flows computed at high-Reynolds-number flight-test conditions show good agreement with surface and offsurface visualizations obtained in flight.
This article describes the numerical simulation of the unsteady viscous flow around the F-18 aircraft at high angles of attack. A generalized overset zonal grid scheme is used to decompose the computational space around the complete aircraft, included deflected control surfaces. The grids around various components of the aircraft are created numerically using a three-dimensional hyperbolic grid generation procedure. The Reynolds-averaged Navier-Stokes equations are integrated using a time-accurate, implicit procedure. Results for the turbulent flow around the F-18 aircraft at 30 deg angle of attack show the details of the flowfield structure, including the unsteadiness created by the vortex burst and the resulting fluctuating airloads exerted on the vertical tail. The computed results agree fairly well with flight data for surface pressure, surface flow pattern, vortex burst location, and the dominant frequency for tail load fluctuations.
Three-dimensional thin-layer Navier-Stokes computations are presented for the F/A-18 configuration. The modeled configuration includes an accurate surface representation of the fuselage, leading-edge extension (LEX), and wing, both with and without leading-edge flap deflection. A multiblock structured grid strategy is employed to decompose the computational flowfield domain around the subject configuration. Steady-state solutions are obtained from an algorithm that solves the compressible Navier-Stokes equations with an upwind-biased, nux-difference splitting approach. The results presented are based on a fully turbulent flow assumption, simulating the high Reynolds number flow conditions that correspond to a recent F/A-18 flight experiment. Good agreements between the computations and the flight test results are obtained for both surface flow patterns as well as surface pressure distributions. Furthermore, a correlation between the computed LEX vortex-core and the flight test results, observed by way of smoke visualization, is also presented.
The generation of significant side forces and yawing moments on an F/A-18 fuselage through tangential slot blowing is analyzed using computational fluid dynamics. The effects of freestream Mach number, jet exit conditions, jet length, and jet location are studied. The effects of over- and underblowing on force and moment production are analyzed. Non-time-accurate solutions are obtained to determine the steady-state side forces, yawing moments, and surface pressure distributions generated by tangential slot blowing. Time-accurate solutions are obtained to study the force onset time lag of tangential slot blowing. Comparison with available experimental data from full-scale wind-tunnel and subscale wind-tunnel tests are made. This computational analysis complements the experimental results and provides a detailed understanding of the effects of tangential slot blowing on the flowfield about the isolated F/A-18 forebody. Additionally, it extends the slot-blowing database to transonic maneuvering Mach numbers.
A full-scale F/A-18 was tested in the 80- by 120-Foot Wind Tunnel at NASA Ames Research Center to measure the effectiveness of pneumatic forebody vortex control devices. By altering the forebody vortex flow, yaw control can be maintained to angles of attack greater than 50 deg. Two forebody vortex control devices were tested: a discrete circular jet and a tangential slot. The tests were conducted for angles of attack between 25 and 50 deg, and angles of sideslip from 0 to +/- 15 deg. The Reynolds number based on wing mean aerodynamic chord ranged from 4.5 x 10 exp 6 to 12.0 x 10 exp 6. The time-averaged side forces and yawing moments, along with both time-averaged and time-dependent pressures on the forebody of the aircraft are presented here for various configurations. Of particular interest was the results that the tangential slot blowing had a greater effect on the yawing moment than the discrete circular jet. Additionally, it was found that blowing very close to the radome apex was not as effective as blowing slightly farther aft on the radome, and that a 16-inch slot was more effective than either an 8- or 48-inch long slot.
Tail buffet studies were conducted on a full-scale, production, F/A-18, fighter aircraft in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center in Moffett Field, California. The F/A-18 was tested over an angle-of-attack range of 18deg to 50deg, a side-slip range of -15deg to 15deg, and at wind speeds of up to 100 knots. The maximum speed corresponds to a Reynolds number of 12.3 x 10(exp 6) based on mean aerodynamic chord and a Mach number of 0.15. The port, vertical tail fin was instrumented with thirty-two surface pressure transducers, arranged in four by four arrays on both sides on the fin. The aircraft was also equipped with a removable Leading Edge eXtension (LEX) fence that is used on F/A-18 aircraft to reduce tail buffet loads. Time-averaged, power-spectral analysis results are presented for the tail fin bending moment derived from the integrated pressure field. The results are only for the zero side-slip condition, both with and without the LEX fence. The LEX fence significantly reduces the magnitude of the root-mean-square pressures and bending moments. Scaling issues are addressed by comparing full-scale results for pressures at the 60%-span and 45%-chord location with published results of small-scale, F/A-18 tail-buffet tests. The comparison shows that the tail buffet frequency scales very well with length and velocity. Root-mean-square pressures and power spectra do not scale as well. The LEX fence is shown to reduce tail buffet loads at all model scales.
The effect of vehicle configuration and flight control system performance on the roll agility of a modern fighter aircraft has been investigated. A batch simulation of a generic F-18 Hornet was used to study the roll agility as measured by the time to roll through 90 deg metric. Problems discussed include definition of agility, factors affecting the agility of a vehicle, the development of the time to roll through 90 deg agility metric, and a simulation experiment. It is concluded that the integral of stability or wind axis roll rate should be used as a measure of the roll measure traversed. The time through roll angle 90 deg metric is considered to be a good metric for measuring the transient performance aspect of agility. Roll agility of the F-18, as measured by 90 deg metric, can be improved by 10 to 30 percent. Compatible roll and rudder actuator rates can significantly affect 90 deg agility metric.
A transonic wind tunnel test of an 8% F/A-18E model was conducted in the NASA Langley Research Center (LaRC) 16-Foot Transonic Tunnel (16-Ft TT) to investigate the Abrupt Wing Stall (AWS) characteristics of this aircraft. During this test, both steady and unsteady measurements of balance loads, wing surface pressures, wing root bending moments, and outer wing accelerations were performed. The test was conducted with a wide range of model configurations and test conditions in an attempt to reproduce behavior indicative of the AWS phenomenon experienced on full-scale aircraft during flight tests. This paper focuses on the analysis of the unsteady data acquired during this test. Though the test apparatus was designed to be effectively rigid. model motions due to sting and balance flexibility were observed during the testing, particularly when the model was operating in the AWS flight regime. Correlation between observed aerodynamic frequencies and model structural frequencies are analyzed and presented. Significant shock motion and separated flow is observed as the aircraft pitches through the AWS region. A shock tracking strategy has been formulated to observe this phenomenon. Using this technique, the range of shock motion is readily determined as the aircraft encounters AWS conditions. Spectral analysis of the shock motion shows the frequencies at which the shock oscillates in the AWS region, and probability density function analysis of the shock location shows the propensity of the shock to take on a bi-stable and even tri-stable character in the AWS flight regime.
Flight and radar position records are analyzed to determine the winds encountered by four airliners that penetrated a microburst on approach to Denver's Stapleton International Airport on July 11, 1988. The four encounters provide information about the time-varying changes in the strength, size, and location of the microburst phenomenon. The results show significant expansion in the size of the microburst and indicate the presence of fluctuations in the internal wind velocity associated with multiple microburst cells. At its peak strength, as experienced by the second aircraft, the microburst produced a headwind-to-tailwind velocity change of 115 ft/s. It is shown that the developing wind patterns derived from the flight-data analysis are in general agreement with results derived from ground-based Doppler weather radar and from a numerical microburst simulation. The data from the four aircraft complement these other findings by providing a more detailed analysis of the microburst's internal wind environment.
On September 8, 1994, a Boeing 737-300 passenger airplane was on a downwind approach to the Pittsburgh International Airport at an altitude of 5000 feet above ground level (6000 feet MSL). While in a shallow left turn onto a downwind approach heading, the airplane crossed into the vortex trail of a Boeing 727 flying in the same approach pattern about 4 miles ahead. The B-737 airplane rolled and turned sharply to the left, exited the vortex wake and plunged into the ground. Weather was not a factor in the accident. The airplane was equipped with a 11+ channel digital Flight Data Recorder (FDR) and a multiple channel Cockpit Voice Recorder (CVR). Both recorders were recovered from the crash site and provided excellent data for the development of an accident scenario. Radar tracking of the two airplanes as well as the indicated air speed (IAS) perturbations clearly visible on the B-737 FDR recordings indicate that the upset was apparently initiated by the airplane's crossing into the wake of the B-727 flying ahead in the same traffic pattern. A 6 degree-of-freedom simulation program for the B-737 airplane using MATLAB and SIMULINK was constructed. The simulation was initialized at the stabilized flight conditions of the airplane about 13 seconds prior to its entry into the vortex trail of the B-727 airplane. By assuming a certain combination of control inputs, it was possible to produce a simulated motion that closely matched that recorded on the FDR.
Various inlet-engine combinations have been studied to find a preferred inlet concept for integration with an advanced technology Mach 2.2 cruise vehicle having a cruise lift-to-drag ratio of 9.6. For the purposes of this study, the range capability for a fixed takeoff gross weight was used to assess the various inlet-engine combinations. Inlet concept selection studies are described which indicated that an axisymmetric, mixed compression inlet was preferred. This study considered four inlet and three engine cycle combinations where the engine airflow was tailored to the inlet airflow delivery capability. Detailed design studies of two mixed compression inlet types are discussed. These were a translating centerbody inlet and a collapsing centerbody bicone inlet. The aerodynamic and mechanical design of each inlet is described. These inlets were also matched to different engine cycles tailored to the inlet airflow capability. The range increments favored the bicone inlet concept primarily because of lighter weight, reduced bleed air, and greater transonic airflow/thrust capability.
An experimental study of the Lissaman 7769 and Miley M06-13-128 airfoils at low chord Reynolds numbers is presented. Although both airfoils perform well near their design Reynolds number of about 600,000, they each produce a different type of hysteresis loop in the lift and drag forces when operated below chord Reynolds numbers of 300,000. The type of hysteresis loop was found to depend upon the relative location of laminar separation and transition. The influence of disturbance environment and experimental procedure on the low Reynolds number airfoil boundary-layer behavior is also presented.
The paper reports on results of heat-transfer tests conducted on a 1/29-scale model of the X-24C-12I hypersonic research aircraft configuration in a Mach 6 tunnel at a Reynolds number of thirteen million using the phase-change heat transfer technique. Sequences of phase-change heat transfer pattern photographs are presented showing windward side and leeward side heating processes. Theoretical predictions of dimensionless heat transfer coefficients along a data line on lower fuselage and on fuselage side bracket the experimental values. A turbulent heating theory gives good agreement with data when shifted to a new virtual origin.
The X-29A advanced technology demonstrator has shown the practicality and advantages of the capability to compute and display, in real-time, aeroperformance flight results. This capability includes the calculations of the in-flight measured drag polar, lift curve, and aircraft specific excess power. From these elements, many other types of aeroperformance measurements can be computed and analyzed. The technique can be used to give an immediate postmaneuver assessment of data quality and maneuver technique, thus, increasing the productivity of a flight program. A key element of this new method was the concurrent development of a real-time, in-flight, net-thrust algorithm, based on the simplified gross thrust method. This net-thrust algorithm allows for the direct calculation of total aircraft drag.
This paper presents a technique for comparing, in real time, the flight-test time histories for X-29A aircraft with time histories computed from linearized mathematical models. Such a comparison allows the flight-test personnel to verify that the aircraft is performing as predicted, to determine regions of nonlinear behavior, and to increase the rate of envelope expansion. The types of mathematical modeling and equipment required, the procedure used, and actual flight-test results are discussed.
An extension of the program STARS (a general-purpose structural analysis program) has been developed; this extension implements a complete aeroservoelastic analysis capability. Previous capabilities included finite-element modeling as well as statics, buckling, vibration, dynamic response, and flutter analyses. This paper presents a description and the formulation of STARS in its current state along with example dynamic, aeroelastic, and aeroservoelastic analyses pertaining to the X-29A aircraft. These examples include vibration analysis results as well as flutter analysis results obtained by the conventional k method and the velocity root-contour solution. Finally, selected open- and closed-loop aeroservoelastic analysis results based on a hybrid formulation are compared to illustrate, using the calculated frequency responses, the interactions of structures, aerodynamics, and flight controls.
Results from wind-tunnel tests of two large-scale models of axisymmetric mixed-compression inlet systems designed for Mach number 3.5 are compared. One inlet incorporated a 'traveling'-bleed system in an effort to achieve maximum transonic engine airflow supply. The other inlet required only a 'fixed'-bleed system, but had 21 percent less transonic airflow supply. The inlet with fixed bleed appears more attractive, if auxiliary airflow systems are used to increase the transonic airflow supply, because it can be 45 percent shorter and would be considerably lighter than the traveling-bleed inlet. In addition, the fixed-bleed inlet offers more operating-control margin at supersonic Mach numbers when the inlet is started. Further, its off-design performance is higher because separation of the flow in the subsonic diffuser can be avoided - something that apparently cannot be done with a traveling-bleed inlet without reducing the transonic airflow supply. Finally, it appears that the management and efficiency of bleed airflow for the fixed-bleed inlet can be improved using analytical methods verified by the tests.
Data and correlations for transition from laminar to turbulent flow on 45- and 60-deg swept cylinders are presented. The data were obtained at Mach 3. 5 in the Pilot Low-Disturbance Wind Tunnel at NASA Langley. Freestream noise levels were varied during the test program from extremely low values that were essentially in the instrument noise range to much higher values approaching those in conventional wind tunnels. The results show that end plates or large trips near the upstream end of the cylinders cause turbulent flow along the entire attachment line of the models over the freestream test Reynolds number range (based on cylinder diameter) of approximately 1. 0 multiplied by 10**5 less than R infinity ,//D less than 1. 6 multiplied by 10**6. When all end disturbance sources are removed, transition occurs on the attachment lines at R// infinity ,//D congruent 7-8 multiplied by 10**5 independent of freestream noise levels and in agreement with previous correlations. With the addition of small roughness elements on the cylinder attachment lines, transition occurs at lower values of the Reynolds number, depending on both the roughness height and the wind-tunnel noise level.
A global model approach developed on data from the experimental aircraft program X-31 EFM was recently improved during the X-31 VECTOR program. Both programs are briefly presented, focusing on extremely precise ESTOL landings following a slow, thrust-vectored approach at high angle of attack. High-accuracy navigation and inertial sensor systems enable onboard calculation of the height above runway with sufficient accuracy which is also required to identify ground effect. System identification procedure utilizes a global model to cover the entire flight regime including high-lift configuration during power approach and landing. Analysis of specially designed flight test led to aerodynamic increment tables for supplementary update of the original database. The wind-tunnel and computational fluid dynamics (CFD) predicted ground effect was incorporated into the higly accurate identified global model, which was then used for conventional landings resimulation to support the initial extremely short takeoff and landing (ESTOL) to the ground flight clearance process. Recently an incremental ground effect model was implemented to supplement the original data set. First identification results from conventional and ESTOL landings show some improvements compared to the predictions at very low height above ground.
Parameter identification of the X-31A experimental aircraft was conducted throughout the envelope expansion using mainly pilot input maneuvers, The quality of the identification results, especially in the high-angle-of-attack regime, suffered from high correlations between the aircraft controls and states as well as from insufficient sideslip excitation. These problems are caused by the high feedback gains used in the flight control system to control the basically unstable airplane, Therefore, in 1994-1995 a dedicated Right test program using single surface excitation for parameter identification at high angles of attack was conducted, The results prove that identification of highly augmented aircraft benefits considerably from this technique, This article presents the realization of the single-surface excitation using the X-31A flutter test box, Also, the Eight test program, the applied evaluation methods, and the identification models are discussed. Selected identification results from the single-surface excitation tests are presented and compared to those obtained from the pilot input maneuvers.
Computations of lift and drag polars for a transport aircraft wing/fuselage high lift configuration using the MEGAFLOW code system are carried out and compared to windtunnel experiments. The main emphasis is laid on a comparison of the block-structured and the unstructured code modules for such type of application. For the block-structured FLOWer code in combination with a k-w turbulence model the numerical results are in good agreement with the available experimental data in the linear CL range. Beyond 150 incidence a strong separation near the flap cut-out is simulated, leading to an underprediction of total lift near CL, max compared to the experimental data. In contrast to this, the results ofthe unstructured TAU code utilizing the Spalart-Allmaras turbulence model are characterized by a nearly constant lift overestimation up to maximum lift without the aforementioned separation tendency at moderate incidences. The lift overprediction in the unstructured results is attributed to the main wing and the slat upperside suction peaks, which are higher resolved by the unstructured grid. Neither code reproduces the lift breakdown beyond CL, max according to the experiments. The use of preconditioning in conjunction with the FLOWer code shows only minor improvement of the accuracy, but considerable deterioration of the convergence properties, requiring improvements for routine use. Further studies will focus on the influence of geometry simplifications at the wing root in the theoretical models and its impact on the experimental evidence.
A series of wind-tunnel and laboratory tests were conducted at the NASA Langley V/STOL tunnel facility to determine both the detailed structure and the induced effects of aspect-ratio-4.0 rectangular jets both in a subsonic crosswind and in quiescent conditions. Wind-tunnel tests were conducted on both blunt (nozzle major axis normal to free stream) and streamwise (nozzle major axis parallel to free stream) nozzle orientations for jet injection angles ranging from 15 to 90 degrees at jet-to-crossflow velocity ratios of 4, 8, and 10. Results indicate that the blunt nozzle induced effects are more significant than those produced by comparable streamwise-oriented jets and that both the flow-field structure and induced effects of streamwise-oriented rectangular jets are quite similar to those created by round jets. Additionally, it is shown that significant differences exist in the vortex flow fields generated by the same rectangular nozzle mounted in two different test hardware configurations.