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NASA NTRSDate not supplied

Turbulence Simulations of Transonic Flows over NACA-0012 and OAT15A Airfoils

Three different simulation approaches, namely unsteady Reynolds-averaged Navier-Stokes (URANS), delayed detached-eddy simulation (DDES), and wall-modeled large-eddy simulation (WMLES) are employed to simulate transonic flow over an NACA-0012, OAT15A airfoils at different angles of attack covering pre- and post-buffet-onset regimes. The freestream Mach number is 0.75, and the Reynolds number based on the chord length is 10 and 3million, respectively. The NASA FUN3D solver is used for the simulations, which is an unstructured, compressible flow solver. The URANS simulations are performed using the Spalart-Allmaras (SA) model with the compressibility correction, the DDES predictions are based on the SA model, and the WMLES are performed using an equilibrium wall-model. The unsteady RANS simulations, only with the compressibility correction, predict the pre- and post- buffet characteristics, which compare well with the experimental results. DDES results predicted a lower buffet onset angle compared to experiment. The predicted shock locations are upstream of the locations predicted by URANS. Using a fine grid in the spanwise direction, WMLES predictions show buffeting consistent with the experiment.

Computational Fluid Dynamics↗
NASA NTRS1991

Dynamic response of a hammerhead launch vehicle wind-tunnel model

NASA-Langley has wind tunnel-tested an aeroelastically-scaled (1/10th-scale) Atlas-Centaur I Large Payload Fairing launch-vehicle model capable of simulating either of the first two bending vibration modes of the full-scale vehicle, on the basis of a partial-mode technique. While the primary emphasis was on the vehicle's buffet response, angle-of-attack studies were conducted for several payload fairing configurations with a view to both the buffet response and the dynamic stability of off-design conditions. No dynamic instabilities were discovered among the range of configurations, and payload fairing L/D variations were found to have only small buffet effects except for the smallest such value, in the second bending mode configuration.

Cole, Stanley R.↗
NASA NTRS1993

Dynamic response characteristics of two transport models tested in the National Transonic Facility

This paper documents recent experiences with measuring the dynamic response characteristics of a commercial transport and a military transport model during full scale Reynolds number tests in the National Transonic Facility. Both models were limited in angle of attack while testing at full scale Reynolds number and cruise Mach number due to pitch or stall buffet response. Roll buffet (wing buzz) was observed for both models at certain Mach numbers while testing at high Reynolds number. Roll buffet was more severe and more repeatable for the military transport model at cruise Mach number. Miniature strain-gage type accelerometers were used for the first time for obtaining dynamic data as a part of the continuing development of miniature dynamic measurements instrumentation for cryogenic applications. This paper presents the results of vibration measurements obtained for both the commercial and military transport models and documents the experience gained in the use of miniature strain gage type accelerometers.

Young, Clarence P., Jr.↗
NASA NTRS1944

Aerodynamic Tests of a Full-scale TBF-1 Aileron Installation in the Langley 16-foot High-Speed Tunnel

The failure of wing panels on a number of TBF-1 and TBM-1 airplanes in flight has prompted several investigations of the possible causes of failure. This report describes tests in the Langley 16-foot high-speed tunnel to determine whether these failures could be attributed to changes in the aerodynamic characteristics of the ailerons at high speeds. The tests were made of a 12-foot-span section including the tip and aileron of the right wing of a TBF-1 airplane. Hinge moments, control-link stresses due to aerodynamic buffeting, and fabric-deflection photographs were obtained at true airspeeds ranging from 110 to 365 miles per hour. The aileron hinge-moment coefficients were found to vary only slightly with airspeed in spite of the large fabric deflections that developed as the speed was increased. An analysis of these results indicated that the resultant hinge moment of the ailerons as installed in the airplane would tend to restore the ailerons to their neutral position for all the high-speed flight conditions covered in the tests. Serious aerodynamic buffeting occurred at up aileron angles of -10 degrees or greater because of stalling of the sharp projecting lip of the Frise aileron. The peak stresses set up in the aileron control linkages in the buffeting condition were as high as three times the mean stress. During the hinge-moment investigation, flutter of the test installation occurred at airspeeds of about 150 miles per hour. This flutter condition was investigated in some detail and slow-motion pictures were made of the motion of the wing tip and aileron. The flutter was found to involve simultaneous normal bending and chordwise oscillation of the wing and flapping of the aileron. The aileron motion appeared to be coupled with this flutter condition and was investigated in some detail and slow-motion pictures were made of the motion of the wing tip and aileron. The flutter was found to involve simultaneous normal bending and chordwise oscillation of the wing and flapping of the aileron. The aileron motion appeared to be coupled with the motion of the wing through the mass unbalance of the aileron in the normal-to-chord plane due to location of the hinge line 2.17 inches below the center of gravity of the aileron. Flutter did not occur when the installation was stiffened to prevent chordwise motion or when the bending frequency of the aileron system was appreciably higher than that of the wing as in the complete airplane installation.

Becker, John V↗
NASA NTRS1947

Wind-Tunnel Tests of a 0.182-Scale Model of an F4U-1 Airplane with External Stores

Tests were made in the Langley 7- by 10-foot tunnel on a 0.182-scale model of an F4U-1 airplane with external stores. This paper is concerned mainly with presenting the data obtained in this investigation and with a comparison of some of these data with flight-test results determining the feasibility of estimating flight buffet Mach number from tunnel data. The results of this investigation indicate that the incremental drag coefficient due to external stores may be used to estimate the maximum Mach number that the F4U-1 airplane may reach in flight when it is equipped with external stores. This estimation is conservative for the five configurations investigated by mounts varying from 0 to 10 percent of the flight limit Mach number. The free-stream tunnel Mach number corresponding to sonic flow over the lower surface of the wing in the region of the store is a good indication of the lower limit of buffet in flight of the F4U-1 airplane when equipped with external stores. The fluctuations of total pressure over the horizontal tail are not sufficiently large (maximum of 1 percent q(sub o) to cause buffeting of the airplane.

Silvers, H. Norman↗
NASA NTRS2015

Initial Assesment of Space Launch System Transonic Unsteady Pressure Environment

A series of wind tunnel tests were conducted at the NASA Langley Research Center Transonic Dynamics Tunnel to assess the transonic buffet environment for the Space Launch System (SLS) launch vehicle. An initial test, conducted in 2012, indicated an elevated buffet environment prompting a second test to provide further insight into the buffet phenomena and assess potential solutions to reduce the response levels of these environments. During the course of the test program, eight variants of the SLS-10000 configuration were examined. The effect of these configuration variants on the coefficient of the root-mean-square fluctuation of pressure about the mean as a function of test condition indicates that the maximum fluctuating pressure levels are extremely sensitive to the geometry of the forward attachment of the solid rocket boosters (SRBs) to the SLS Core. The addition of flow fences or changes to the SRB nose cone geometry can alleviate the unsteady pressure environment.

Sekula, Martin K.↗
NASA NTRS2017

Inverse Force Determination on a Small Scale Launch Vehicle Model Using a Dynamic Balance

A launch vehicle can experience large unsteady aerodynamic forces in the transonic regime that, while usually only lasting for tens of seconds during launch, could be devastating if structural components and electronic hardware are not designed to account for them. These aerodynamic loads are difficult to experimentally measure and even harder to computationally estimate. The current method for estimating buffet loads is through the use of a few hundred unsteady pressure transducers and wind tunnel test. Even with a large number of point measurements, the computed integrated load is not an accurate enough representation of the total load caused by buffeting. This paper discusses an attempt at using a dynamic balance to experimentally determine buffet loads on a generic scale hammer head launch vehicle model tested at NASA Ames Research Center's 11' x 11' transonic wind tunnel. To use a dynamic balance, the structural characteristics of the model needed to be identified so that the natural modal response could be and removed from the aerodynamic forces. A finite element model was created on a simplified version of the model to evaluate the natural modes of the balance flexures, assist in model design, and to compare to experimental data. Several modal tests were conducted on the model in two different configurations to check for non-linearity, and to estimate the dynamic characteristics of the model. The experimental results were used in an inverse force determination technique with a psuedo inverse frequency response function. Due to the non linearity, the model not being axisymmetric, and inconsistent data between the two shake tests from different mounting configuration, it was difficult to create a frequency response matrix that satisfied all input and output conditions for wind tunnel configuration to accurately predict unsteady aerodynamic loads.

testing↗
NASA NTRS2018

Multidisciplinary Design Optimization and Cruise Mach Number Study of Truss-Braced Wing Aircraft

The Subsonic Ultra Green Aircraft Research (SUGAR) Phase III was led by Dr. Rakesh K. Kapania and Dr. Joseph A. Schetz at the Multidisciplinary Analysis and Design Center for Advanced Vehicles, Department of Aerospace and Ocean Engineering, Virginia Tech, Blacksburg VA. The research was performed from December 2014 to December 2015. Three major areas were investigated: Multidisciplinary Design Optimization (MDO) studies of truss braced wing (TBW) and strut braced wing (SBW) vehicles at cruise Mach numbers of 0.7 and 0.8 for a flight mission similar to current market single aisle configurations. The performance and the characteristics of the optimized vehicles were compared to the SUGAR Phase II TBW vehicle. These results were obtained without applying any of the extended transonic aerodynamic and aeroelastic tools that will be discussed later. It was found that the cruise Mach number has a large effect on the best “truss” configuration. At Mach 0.7, an SBW has a better fuel consumption and better take-off gross weight (TOGW). However, at Mach 0.8, the TBW is superior because the jury strut aids in satisfying the flutter constraint; Two-dimensional, steady, transonic aerodynamic analysis of the Boeing Airfoil J (BACJ) airfoil was performed for a range of thickness ratios, Mach numbers and lift coefficients. Reynolds-averaged Navier-Stokes (RANS) equations were solved to obtain the lift-curve slope, wave drag coefficient, the location of the center of pressure and to predict the separation at the trailing edge, which may lead to buffeting. One of the goals was to develop a database of lift-curve slope and the location of center of pressure, which could be used in a transonic aeroelastic analysis. Another goal was to compare the wave drag coefficients to those predicted by Lock’s fourth-power law and also to compare the transonic effects obtained from RANS simulations to those predicted by the Korn equations. A third goal was to develop a buffet boundary, which can be integrated into the MDO framework to prevent the optimized designs from probable buffeting; A state-space transonic aeroelastic analysis tool was developed, which can incorporate the nonlinear transonic effects in the unsteady aerodynamics but is yet computationally cheap when used within the VT MDO framework. The aeroelastic analysis uses Leishman- Beddoes (LB) indicial functions, which generated a state-space representation of the aeroelastic system. The indicial functions allow the incorporation of data for steady lift-curve slope and location of the center of pressure. Thus, the steady transonic effects are included, and the unsteady aerodynamic responses are a linearization about the steady results. The aeroelastic approach discretizes the wing into numerous strips, which results in a large eigenvalue problem as each strip has eight augmented aerodynamic states as per the LB theory. Thus, to reduce the computation expense, a reduced order model (ROM) was developed. The approach was validated using a few examples.

Kapania, Rakesh K.↗
NASA NTRSDate not supplied

Comparison of Corcos-based and experimentally-derived coherence factors for BFFs estimation

In this paper, high-spatial-resolution unsteady Pressure Sensitive Paint (uPSP) data are utilized to compare two methods for panel Buffet Forcing Functions (BFF) estimation for the Space Launch System (SLS). Such methods are based on discrete pressure measurements within a panel but employ coherence factors to account for partially-correlated fluctuating pressures across the whole panel. In one method, coherence factors are derived based on the Corcos model, whereas the second method utilizes experimentally-derived coherence factors. To simulate discrete measurements using uPSP data, suitable subsets of the data are extracted. When full uPSP resolution is retained, uPSP data provide a benchmark to assess discrete-measurements-based methods. The analysis focuses on the peak SLS buffet environment located downstream of the Forward Attachment Hardware (FAH) between the core stage and solid rocket boosters. Trends of Corcos-based and experimentally-derived coherence factors are in reasonable agreement with the benchmark. However, at certain frequencies, experimentally-derived coherence factors are sensitive to the separation distance between pressure measurements utilized to compute coherence lengths. Such sensitivity originates from deviation of the experimental-based coherence function from an exponential decay assumption. On the other hand, the present implementation of the Corcos model fails to capture certain non-turbulent boundary layer related environments, such as a subharmonic of FAH vortex-shedding. For all methods presented in this paper, at near transonic conditions, increased pressure coherence and spatial nonuniformity lead to BFF overestimation and sensitivity to the pressure measurement location within the panel.

transonic buffet↗
NASA NTRSDate not supplied

Simulations of Mach 0.8 Transonic Truss-Braced Wing Aircraft Aerodynamics at High Angles of Attack

This paper presents numerical simulations of the Mach 0.8 Transonic Truss-Braced Wing (TTBW) aircraft at high angle of attack using CFD solver FUN3D. Three different approaches - steady Reynolds Averaged Navier-Stokes (RANS), Unsteady Reynolds Averaged Navier-Stokes (URANS), and Delayed Detached Eddy Simulation (DDES) are used to simulation the TTBW aircraft at Mach number 0.8 and altitude 40, 000 ft. The pitch break phenomenon is observed at this flight condition. The simulation results show that the pitch break angle of attack occurs at 2.75˚ by DDES, 2.81◦ by URANS, and 2.89˚ by RANS. The DDES unsteady pressure begins to grow at angle of attack of 3.5˚, which might indicate the buffet onset, and rises one order of magnitude larger at angle of attack of 4.0˚. However, the URANS unsteady pressure begins to grow at angle of attack of 4.0˚. Both unsteady simulation results show the pitch break occurs before buffet onset.

TTBW↗
NASA NTRS1973

Dynamic effects of shock-induced flow separation.

Shock-induced flow separation is the flow mechanism usually responsible for what the structural dynamicist terms 'buffet.' The shock-induced flow separation affects the aeroelastic response via two different mechanisms: (1) the flow separation generates fluctuating pressures, i.e., a forcing function that is independent of the motion of the aerodynamic surface, e.g., an aircraft wing, and (2) the flow separation affects the motion-dependent forces and can in some cases generate negative aerodynamic damping. A simple analysis is presented which, using static experimental data as an input, can predict these two buffet-components for a wing in high Mach number subsonic flow.

Ericsson, L. E.↗
NASA NTRS1977

Separated-flow unsteady pressures and forces on elastically responding structures

Broadband rms, spectral density, and spatial correlation information that characterizes the fluctuating pressures and forces that cause aircraft buffet is presented. The main theme of the paper in describing buffet excitation is to show the effects of elasticity. Data are presented that were obtained in regions of separated flow on wings of wind-tunnel models of varying stiffness and on the wing of a full scale aircraft. Reynolds number effects on the pressure fluctuations are also discussed.

Coe, C. F.↗
NASA NTRS1978

Delta method, an empirical drag buildup technique

An empirical drag correlation technique was developed from analysis of 19 subsonic and supersonic military aircraft and 15 advanced or supercritical airfoil configurations which can be applied in conceptual and advanced aircraft design activities. The Delta Method may be used for estimating the clean wing drag polar for cruise and maneuver conditions up to buffet onset, and to approximately Mach 2.0. This technique incorporates a unique capability of predicting the off-design performance of advanced or supercritical airfoil sections. The buffet onset limit may also be estimated. The method is applicable to wind tunnel models as well as to full scale configurations. This technique has been converted into a computer code for use on the IBM 360 and CDC 7600 computer facilities at NASA AMES. Results obtained using this method to predict known aircraft characteristics are good and agreement can be obtained within a degree of accuracy judged to be sufficient for the initial processes of preliminary design.

Feagin, R. C.↗
NASA NTRS1981

Historical background and design evolution of the transonic aircraft technology supercritical wing

Two dimensional wind tunnel test results obtained for supercritical airfoils indicated that substantial improvements in aircraft performance at high subsonic speeds could be achieved by shaping the airfoil to improve the supercritical flow above the upper surface. Significant increases in the drag divergence Mach number, the maximum lift coefficient for buffer onset, and the Mach number for buffet onset at a given lift coefficient were demonstrated for the supercritical airfoil, as compared with a NACA 6 series airfoil of comparable thickness. These trends were corroborated by results from three dimensional wind tunnel and flight tests. Because these indicated extensions of the buffet boundaries could provide significant improvements in the maneuverability of a fighter airplane, an exploratory wind tunnel investigation was initiated which demonstrated that significant aerodynamic improvements could be achieved from the direct substitution of a supercritical airfoil on a variable wing sweep multimission airplane model.

Ayers, T. G.↗
NASA NTRS1974

Wind-tunnel investigation of aerodynamic load distribution on a variable-wing-sweep fighter airplane with a NASA supercritical airfoil

Wind-tunnel tests have been conducted at Mach numbers of 0.85, 0.88, and 0.90 to determine the aerodynamic load distribution for the 39 deg swept-wing configuration of a variable-wing-sweep fighter airplane with a NASA supercritical airfoil. Chordwise pressure distributions were measured at two wing stations. Also measured were the overall longitudinal aerodynamic force and moment characteristics and the buffet characteristics. The analysis indicates that localized regions of shock-induced flow separation may exist on the rearward portions of the supercritical wing at high subsonic speeds, and caution must be exercised in the prediction of buffet onset when using variations in trailing-edge pressure coefficients at isolated locations.

Hallissy, J. B.↗
NASA NTRS1982

Comparison of Flight Measured, Predicted and Wind Tunnel Measured Winglet Characteristics on a KC-135 Aircraft

One of the objectives of the KC-135 Winglet Flight Research and Demonstration Program was to obtain experimental flight test data to verify the theoretical and wind tunnel winglet aerodynamic performance prediction methods. Good agreement between analytic, wind tunnel and flight test performance was obtained when the known differences between the tests and analyses were accounted for. The flight test measured fuel mileage improvements for a 0.78 Mach number was 3.1 percent at 8 x 10(5) pounds W/delta and 5.5 percent at 1.05 x 10(6) pounds W/delta. Correcting the flight measured data for surface pressure differences between wind tunnel and flight resulted in a fuel mileage improvement of 4.4 percent at 8 x 10(5) pounds W/delta and 7.2 percent at 1.05 x 10(6) pounds W/delta. The performance improvement obtained was within the wind tunnel test data obtained from two different wind tunnel models. The buffet boundary data obtained for the baseline configuration was in good agreement with previous established data. Buffet data for the 15 deg cant/-4 deg incidence configuration showed a slight improvement, while the 15 deg cant/-2 deg incidence and 0 deg cant/-4 deg incidence data showed a slight deterioration.

Dodson, R. O., Jr.↗
NASA NTRS1983

DC-10 winglet flight evaluation

The results of a flight evaluation of winglets on a DC-10 Series 10 aircraft are presented. For sensitive areas of comparison, effects of winglets were determined back to back with and without winglets. Basic and reduced span winglet configurations were tested. After initial encounter with low speed buffet, a number of acceptable configurations were developed. For maximum drag reduction at both cruise and low speeds, lower winglets were required, having leading edge devices on upper and lower winglets for the latter regime. The cruise benefits were enhanced by adding outboard aileron droop to the reduced span winglet aircraft. Winglets had no significant impact on stall speeds, high speed buffet boundary, and stability and control characteristics. Flutter test results agreed with predictions and ground vibration data. Flight loads measurement also agreed with predictions.

Source record↗
NASA NTRS1983

DC-10 winglet flight evaluation

Results of a flight evaluation of winglets on a DC-10 Series 10 aircraft are presented. For sensitive areas of comparison, effects of winglets were determined back-to-back with and without winglets. Basic and reduced-span winglet configurations were tested. After initial encounter with low-speed buffet, a number of acceptable configurations were developed. For maximum drag reduction at both cruise and low speeds, lower winglets were required, having leading edge devices on upper and lower winglets for the latter regime. The cruise benefits were enhanced by adding outboard aileron droop to the reduced-span winglet aircraft. Winglets had no significant impact on stall speeds, high-speed buffet boundary, and stability and control. Flutter test results agreed with predictions and ground vibration data. Flight loads measurement, provided in a concurrent program, also agreed with predictions. It was estimated that a production version of the aircraft, using the reduced-span winglet and aileron droop, would yield a 3-percent reduction in fuel burned with capacity payload. This range was 2% greater than with winglets. A 5% reduction in takeoff distance at maximum takeoff weight would also result.

Taylor, A. B.↗