Performance of Swashplateless Ultralight Helicopter Rotor with Trailing-edge Flaps for Primary Flight Control

NASA Technical Reports Server (NTRS)

Shen, Jin-Wei; Chopra, Inderjit

2003-01-01

The objective of present study is to evaluate the rotor performance, trailing-edge deflections and actuation requirement of a helicopter rotor with trailing-edge flap system for primary flight control. The swashplateless design is implemented by modifying a two-bladed teetering rotor of an production ultralight helicopter through the use of plain flaps on the blades, and by replacing the pitch link to fixed system control system assembly with a root spring. A comprehensive rotorcraft analysis based on UMARC is carried out to obtain the results for both the swashplateless and a conventional baseline rotor configuration. The predictions show swashplateless configuration achieve superior performance than the conventional rotor attributed from reduction of parasite drag by eliminating swashplate mechanic system. It is indicated that optimal selection of blade pitch index angle, flap location, length, and chord ratio reduces flap deflections and actuation requirements, however, has virtually no effect on rotor performance.

Wind-tunnel Investigation of High-lift and Stall-control Devices on a 37 Degree Sweptback Wing of Aspect Ratio 6 at High Reynolds Numbers

NASA Technical Reports Server (NTRS)

Koven, William; Graham, Robert R

1948-01-01

Results are presented of an investigation in the Langley 19-foot pressure tunnel of the longitudinal characteristics of a semispan model wing having 37 degrees sweepback of the leading edge, an aspect ratio of 6, and NACA 641-212 airfoil section perpendicular to the 27-percent-chord line. Several types of stall-control devices including extensible round-nose leading-edge flaps, a leading-edge slat, and a drooped leading edge were investigated; partial- and full-span trailing-edge split and double slotted flaps were also tested. In addition, various combinations of the aforementioned leading- and trailing-edge flaps were investigated. The tests covered a range of Reynolds numbers between 2.00 x 10(6) and 9.35 x 10(6). The wing with or without trailing-edge splity of double slotted flap was longitudinally unstable near maximum lift due to tip stalling. The addition of an outboard half-span leading-edge flap or a leading-edge slat to the plain wing or wing with inboard half-span split flaps eliminated tip stalling and resulted in stable moment variations at the stall. The drooped leading edge, on the other hand, was only effective when used in conjunction with an upper-surface fence. The combination of an outboard leading-edge device and inboard half-span double slotted flap resulted in an undesirable loop in the pitching-moment curve near maximum lift in spite of an inboard stall. The loop is attributed to the section characteristics of the double slotted flap. Air-flow surveys behind the wing indicated that a suitably placed horizontal tail would eliminate the loop in the moment curve.

Robust design of multiple trailing edge flaps for helicopter vibration reduction: A multi-objective bat algorithm approach

NASA Astrophysics Data System (ADS)

Mallick, Rajnish; Ganguli, Ranjan; Seetharama Bhat, M.

2015-09-01

The objective of this study is to determine an optimal trailing edge flap configuration and flap location to achieve minimum hub vibration levels and flap actuation power simultaneously. An aeroelastic analysis of a soft in-plane four-bladed rotor is performed in conjunction with optimal control. A second-order polynomial response surface based on an orthogonal array (OA) with 3-level design describes both the objectives adequately. Two new orthogonal arrays called MGB2P-OA and MGB4P-OA are proposed to generate nonlinear response surfaces with all interaction terms for two and four parameters, respectively. A multi-objective bat algorithm (MOBA) approach is used to obtain the optimal design point for the mutually conflicting objectives. MOBA is a recently developed nature-inspired metaheuristic optimization algorithm that is based on the echolocation behaviour of bats. It is found that MOBA inspired Pareto optimal trailing edge flap design reduces vibration levels by 73% and flap actuation power by 27% in comparison with the baseline design.

Effects of flaps on buffet characteristics and wind-rock onset of an F-8C airplane at subsonic and transonic speeds

NASA Technical Reports Server (NTRS)

Monaghan, R. C.; Friend, E. L.

1973-01-01

Wind-up-turn maneuvers were performed to establish the values of airplane normal force coefficient for buffet onset, wing-rock onset, and buffet loads with various combinations of leading- and trailing-edge flap deflections. Data were gathered at both subsonic and transonic speeds covering a range from Mach 0.64 to Mach 0.92. Buffet onset and buffet loads were obtained from wingtip acceleration and wing-root bending-moment data, and wing-rock onset was obtained from airplane roll rate data. Buffet onset, wing-rock onset, and buffet loads were similarly affected by the various combinations of leading- and training-edge flaps. Subsonically, the 12 deg leading-edge-flap and trailing-edge-flap combination was most effective in delaying buffet onset, wing-rock onset, and equivalent values of buffet loads to a higher value of airplane normal force coefficient. This was the maximum flap deflection investigated. Transonically, however, the optimum leading-edge flap position was generally less than 12 deg.

Approach for Structurally Clearing an Adaptive Compliant Trailing Edge Flap for Flight

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Lokos, William A.; Cruz, Josue; Crampton, Glen; Stephens, Craig A.; Kota, Sridhar; Ervin, Gregory; Flick, Pete

2015-01-01

The Adaptive Compliant Trailing Edge (ACTE) flap was flown on the NASA Gulfstream GIII test bed at the NASA Armstrong Flight Research Center. This smoothly curving flap replaced the existing Fowler flaps creating a seamless control surface. This compliant structure, developed by FlexSys Inc. in partnership with Air Force Research Laboratory, supported NASA objectives for airframe structural noise reduction, aerodynamic efficiency, and wing weight reduction through gust load alleviation. A thorough structures airworthiness approach was developed to move this project safely to flight.

Effect of leading- and trailing-edge flaps on clipped delta wings with and without wing camber at supersonic speeds

NASA Technical Reports Server (NTRS)

Hernandez, Gloria; Wood, Richard M.; Covell, Peter F.

1994-01-01

An experimental investigation of the aerodynamic characteristics of thin, moderately swept fighter wings has been conducted to evaluate the effect of camber and twist on the effectiveness of leading- and trailing-edge flaps at supersonic speeds in the Langley Unitary Plan Wind Tunnel. The study geometry consisted of a generic fuselage with camber typical of advanced fighter designs without inlets, canopy, or vertical tail. The model was tested with two wing configurations an uncambered (flat) wing and a cambered and twisted wing. Each wing had an identical clipped delta planform with an inboard leading edge swept back 65 deg and an outboard leading edge swept back 50 deg. The trailing edge was swept forward 25 deg. The leading-edge flaps were deflected 4 deg to 15 deg, and the trailing-edge flaps were deflected from -30 deg to 10 deg. Longitudinal force and moment data were obtained at Mach numbers of 1.60, 1.80, 2.00, and 2.16 for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.16 x 10(exp 6) per foot and for an angle-of-attack range 4 deg to 20 deg at a Reynolds number of 2.0 x 10(exp 6) per foot. Vapor screen, tuft, and oil flow visualization data are also included.

Computational Evaluation of the Steady and Pulsed Jet Effects on the Performance of a Circulation Control Wing Section

NASA Technical Reports Server (NTRS)

Liu, Yi; Sankar, Lakshmi N.; Englar, Robert; Ahuja, K.; Gaeta, R.

2003-01-01

Circulation Control Wing (CCW) technology is a very effective way of achieving very high lift coefficients needed by aircraft during take-off and landing. This technology can also be used to directly control the flow field over the wing. Compared to a conventional high-lift system, a Circulation Control Wing (CCW) can generate the required values of lift coefficient C(sub L,max) during take-off/landing with fewer or no moving parts and much less complexity. Earlier designs of CCW configurations used airfoils with a large radius rounded trailing edge to maximize the lift benefit. However, these designs also produced very high drag. These high drag levels associated with the blunt, large radius trailing edge can be prohibitive under cruise conditions when Circulation Control is no longer necessary. To overcome this difficulty, an advanced CCW section, i.e., a circulation hinged flap was developed to replace the original rounded trailing edge CC airfoil. This concept developed by Englar is shown. The upper surface of the CCW flap is a large-radius arc surface, but the lower surface of the flap is flat. The flap could be deflected from 0 degrees to 90 degrees. When an aircraft takes-off or lands, the flap is deflected as in a conventional high lift system. Then this large radius on the upper surface produces a large jet turning angle, leading to high lift. When the aircraft is in cruise, the flap is retracted and a conventional sharp trailing edge shape results, greatly reducing the drag. This kind of flap does have some moving elements that increase the weight and complexity over an earlier CCW design. But overall, the hinged flap design still maintains most of the Circulation Control high lift advantages, while greatly reducing the drag in cruising condition associated with the rounded trailing edge CCW design. In the present work, an unsteady three-dimensional Navier-Stokes analysis procedure has been developed and applied to this advanced CCW configuration. The solver can be used in both a 2-D and a 3-D mode, and can thus model airfoils as well as finite wings. The jet slot location, slot height, and the flap angle can all be varied easily and individually in the grid generator and the flow solver. Steady jets, pulsed jets, the leading edge and trailing edge blowing can all be studied with this solver.

Force and moment measurements on a 74 deg delta wing with an apex flap

NASA Technical Reports Server (NTRS)

Buter, T. A.; Rao, D. M.

1984-01-01

Results are presented of a subsonic experimental investigation of an apex flap concept on a 74 deg swept delta wing with trailing-edge flaps. The apex flap comprised approximately 6 percent of the wing area forward of a transverse hinge, allowing for upward and downward deflection angles from +40 deg to -20 deg. Upward deflection forces leading-edge vortex formation on the apex flap, resulting in an increased lift component on the apex area. The associated nose-up moment balances the nose-down moment due to trailing-edge flaps, resulting in sizeable increase in the trimmed lift coefficient particularly at low angles of attack. Nose-down apex deflection may be used to augment the pitch control for rapid recovery from high-alpha maneuvers. This report presents the balance data without analysis.

Design of Supersonic Transport Flap Systems for Thrust Recovery at Subsonic Speeds

NASA Technical Reports Server (NTRS)

Mann, Michael J.; Carlson, Harry W.; Domack, Christopher S.

1999-01-01

A study of the subsonic aerodynamics of hinged flap systems for supersonic cruise commercial aircraft has been conducted using linear attached-flow theory that has been modified to include an estimate of attainable leading edge thrust and an approximate representation of vortex forces. Comparisons of theoretical predictions with experimental results show that the theory gives a reasonably good and generally conservative estimate of the performance of an efficient flap system and provides a good estimate of the leading and trailing-edge deflection angles necessary for optimum performance. A substantial reduction in the area of the inboard region of the leading edge flap has only a minor effect on the performance and the optimum deflection angles. Changes in the size of the outboard leading-edge flap show that performance is greatest when this flap has a chord equal to approximately 30 percent of the wing chord. A study was also made of the performance of various combinations of individual leading and trailing-edge flaps, and the results show that aerodynamic efficiencies as high as 85 percent of full suction are predicted.

Optimal placement of trailing-edge flaps for helicopter vibration reduction using response surface methods

NASA Astrophysics Data System (ADS)

Viswamurthy, S. R.; Ganguli, Ranjan

2007-03-01

This study aims to determine optimal locations of dual trailing-edge flaps to achieve minimum hub vibration levels in a helicopter, while incurring low penalty in terms of required trailing-edge flap control power. An aeroelastic analysis based on finite elements in space and time is used in conjunction with an optimal control algorithm to determine the flap time history for vibration minimization. The reduced hub vibration levels and required flap control power (due to flap motion) are the two objectives considered in this study and the flap locations along the blade are the design variables. It is found that second order polynomial response surfaces based on the central composite design of the theory of design of experiments describe both objectives adequately. Numerical studies for a four-bladed hingeless rotor show that both objectives are more sensitive to outboard flap location compared to the inboard flap location by an order of magnitude. Optimization results show a disjoint Pareto surface between the two objectives. Two interesting design points are obtained. The first design gives 77 percent vibration reduction from baseline conditions (no flap motion) with a 7 percent increase in flap power compared to the initial design. The second design yields 70 percent reduction in hub vibration with a 27 percent reduction in flap power from the initial design.

The Effectiveness at High Speeds of a 20-Percent-Chord Plain Trailing-Edge Flap on the NACA 65-210 Airfoil

NASA Technical Reports Server (NTRS)

Stivers, Louis S., Jr.

1947-01-01

An analysis has been made of the lift control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flap on the NACA 65-210 airfoil section. The analysis indicates that the plain trailing-edge flap employed on the 10-percent-thick airfoil at Mach numbers as high as 0.875 retains at least 50-percent of its low-speed lift-control effectiveness, and is sufficiently effective in lateral control application, assuming a rigid wing, to provide adequate airplane rolling characteristics. The plain trailing-edge flap, as compared to the spoiler and the dive-recovery flap, appears to afford the most favorable characteristics as a device for controlling lift continuously throughout the range of Mach numbers from 0.3 to 0.875. At Mach numbers above those for lift divergence of the wing, either a plain flap or a dive-recovery flap may be used on a thin airplane wing to provide auxiliary wing lift when the airplane is to be controlled in flight, other than in dives, at these Mach numbers. The choice of a lift-control device for this use, however, should include the consideration of other factors such as the increments of drag and pitching moment accompanying the use of the device, and the structural and high-speed aerodynamic characteristics of the airplane which is to employ the device.

Wing pressure distributions from subsonic tests of a high-wing transport model. [in the Langley 14- by 22-Foot Subsonic Wind Tunnel

NASA Technical Reports Server (NTRS)

Applin, Zachary T.; Gentry, Garl L., Jr.; Takallu, M. A.

1995-01-01

A wind tunnel investigation was conducted on a generic, high-wing transport model in the Langley 14- by 22-Foot Subsonic Tunnel. This report contains pressure data that document effects of various model configurations and free-stream conditions on wing pressure distributions. The untwisted wing incorporated a full-span, leading-edge Krueger flap and a part-span, double-slotted trailing-edge flap system. The trailing-edge flap was tested at four different deflection angles (20 deg, 30 deg, 40 deg, and 60 deg). Four wing configurations were tested: cruise, flaps only, Krueger flap only, and high lift (Krueger flap and flaps deployed). Tests were conducted at free-stream dynamic pressures of 20 psf to 60 psf with corresponding chord Reynolds numbers of 1.22 x 10(exp 6) to 2.11 x 10(exp 6) and Mach numbers of 0.12 to 0.20. The angles of attack presented range from 0 deg to 20 deg and were determined by wing configuration. The angle of sideslip ranged from minus 20 deg to 20 deg. In general, pressure distributions were relatively insensitive to free-stream speed with exceptions primarily at high angles of attack or high flap deflections. Increasing trailing-edge Krueger flap significantly reduced peak suction pressures and steep gradients on the wing at high angles of attack. Installation of the empennage had no effect on wing pressure distributions. Unpowered engine nacelles reduced suction pressures on the wing and the flaps.

Noise reduction tests of large-scale-model externally blown flap using trailing-edge blowing and partial flap slot covering. [jet aircraft noise reduction

NASA Technical Reports Server (NTRS)

Mckinzie, D. J., Jr.; Burns, R. J.; Wagner, J. M.

1976-01-01

Noise data were obtained with a large-scale cold-flow model of a two-flap, under-the-wing, externally blown flap proposed for use on future STOL aircraft. The noise suppression effectiveness of locating a slot conical nozzle at the trailing edge of the second flap and of applying partial covers to the slots between the wing and flaps was evaluated. Overall-sound-pressure-level reductions of 5 db occurred below the wing in the flyover plane. Existing models of several noise sources were applied to the test results. The resulting analytical relation compares favorably with the test data. The noise source mechanisms were analyzed and are discussed.

The Effectiveness at High Speeds of a 20-Percent-chord Plain Trailing-edge Flap on the NACA 65-210 Airfoil Section

NASA Technical Reports Server (NTRS)

Stivers, Louis S., Jr.

1947-01-01

An analysis has been made of the lift-control effectiveness of a 20-percent-chord plain trailing-edge flap on the NACA 65-210 airfoil section from section lift-coefficient data obtained at Mach numbers from 0.3 to 0.875. In addition, the effectiveness of the plain flap as a lift-control device has been compared with the corresponding effectiveness of both a spoiler and a dive-recovery flag on the INCA 65-210 airfoil section.

Active Flow Separation Control on a NACA 0015 Wing Using Fluidic Actuators

NASA Technical Reports Server (NTRS)

Melton, Latunia P.

2014-01-01

Results are presented from a recent set of wind tunnel experiments using sweeping jet actuators to control ow separation on the 30% chord trailing edge ap of a 30 deg. swept wing model with an aspect ratio (AR) of 4.35. Two sweeping jet actuator locations were examined, one on the flap shoulder and one on the trailing edge flap. The parameters that were varied included actuator momentum, freestream velocity, and trailing edge flap deflection (Delta f ) angle. The primary focus of this set of experiments was to determine the mass flow and momentum requirements for controlling separation on the flap, especially at large flap deflection angles which would be characteristic of a high lift system. Surface pressure data, force and moment data, and stereoscopic particle image velocimetry (PIV) data were acquired to evaluate the performance benefits due to applying active flow control. Improvements in lift over the majority of the wing span were obtained using sweeping jet actuator control. High momentum coefficient, Cu, levels were needed when using the actuators on the ap because they were located downstream of separation. Actuators on the flap shoulder performed slightly better but actuator size, orientation, and spacing still need to be optimized.

Noise Radiation from a Continuous Mold-Line Link Flap Configuration

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Brooks, Thomas F.; Humphreys, William M., Jr.

2011-01-01

The results of an experimental study of the noise from a Continuous Mold-Line Link (CML) flap are presented. Acoustic and unsteady surface pressure measurements were performed on a main element wing section with a half-span CML flap in NASA Langley s Quiet Flow Facility. The acoustic data were acquired with a medium aperture directional array (MADA) of microphones. The Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) method is applied to determine the spatial distribution and strength of the noise sources over the surface of the test model. A Coherent Output Power (COP) method which relates the output from unsteady surface pressure sensors to the output of the MADA is also used to obtain more detailed characteristics of the noise source distribution in the trailing edge region of the CML. These results are compared to those obtained for a blunt flap to quantify the level of noise benefit that is achieved with the CML flap. The results indicate that the noise from the CML region of the flap is 5 to 17 dB lower (depending on flap deflection and Mach number) than the noise from the side edge region of the blunt flap. Lower noise levels are obtained for all frequencies. Spectral analysis of the noise from the cove region of the CML and blunt flap models also reveal a spectral peak in the high frequency range that is related to noise scattering at the trailing edge of the main element. The peaks in the CML and blunt flap cove noise spectra are close in level and often exceed blunt side edge noise. Applying a strip of serrated tape to the trailing edge of the CML flap model main airfoil reduced the peak but increased other noise somewhat. Directivity measurements show that the CML flap can be more directional than the blunt flap.

Noise Radiation from a Continuous Mold-Line Link Flap Configuration

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Brooks, Thomas F.; Humphreys, William M.

2008-01-01

The results of an experimental study of the noise from a Continuous Mold-Line Link (CML) flap are presented. Acoustic and unsteady surface pressure measurements were performed on a main element wing section with a half-span CML flap in NASA Langley s Quiet Flow Facility. The acoustic data were acquired with a medium aperture directional array (MADA) of microphones. The Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) method is applied to determine the spatial distribution and strength of the noise sources over the surface of the test model. A Coherent Output Power (COP) method which relates the output from unsteady surface pressure sensors to the output of the MADA is also used to obtain more detailed characteristics of the noise source distribution in the trailing edge region of the CML. These results are compared to those obtained for a blunt flap to quantify the level of noise benefit that is achieved with the CML flap. The results indicate that the noise from the CML region of the flap is 5 to 17 dB lower (depending on flap deflection and Mach number) than the noise from the side edge region of the blunt flap. Lower noise levels are obtained for all frequencies. Spectral analysis of the noise from the cove region of the CML and blunt flap models also reveal a spectral peak in the high frequency range that is related to noise scattering at the trailing edge of the main element. The peaks in the CML and blunt flap cove noise spectra are close in level and often exceed blunt side edge noise. Applying a strip of serrated tape to the trailing edge of the CML flap model main airfoil, reduced the peak but increased other noise somewhat. Directivity measurements show that the CML flap can be more directional than the blunt flap.

An investigation of unsteady 3D effects on trailing edge flaps

NASA Astrophysics Data System (ADS)

Jost, E.; Fischer, A.; Lutz, T.; Krämer, E.

2016-09-01

The present study investigates the impact of unsteady and viscous three-dimensional aerodynamic effects on a wind turbine blade with trailing edge flap by means of CFD. Harmonic oscillations are simulated on the DTU 10 MW rotor with a flap of 10% chord extent ranging from 70% to 80% blade radius. The deflection frequency is varied in the range between 1p and 6p. To quantify 3D effects, rotor simulations are compared to 2D airfoil computations. A significant influence of trailing and shed vortex structures has been found which leads to a reduction of the lift amplitude and hysteresis effects in the lift response with regard to the flap deflection. In the 3D rotor results greater amplitude reductions and less hystereses have been found compared to the 2D airfoil simulations.

A low speed wind tunnel investigation of Reynolds number effects on a 60-deg swept wing configuration with leading and trailing edge flaps

NASA Technical Reports Server (NTRS)

Rao, Dhanvada M.; Hoffler, Keith D.

1988-01-01

A low-speed wind tunnel test was performed to investigate Reynolds number effects on the aerodynamic characteristics of a supersonic cruise wing concept model with a 60-deg swept wing incorporating leading-edge and trailing-edge flap deflections. The Reynolds number ranged from 0.3 to 1.6 x 10 to the 6th, and corresponding Mach numbers from .05 to 0.3. The objective was to define a threshold Reynolds number above which the flap aerodynamics basically remained unchanged, and also to generate a data base useful for validating theoretical predictions for the Reynolds number effects on flap performance. This report documents the test procedures used and the basic data acquired in the investigation.

Strain Gage Load Calibration of the Wing Interface Fittings for the Adaptive Compliant Trailing Edge Flap Flight Test

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Holguin, Andrew C.; Cruz, Josue; Lokos, William A.

2014-01-01

This is the presentation to follow conference paper of the same name. The adaptive compliant trailing edge (ACTE) flap experiment safety of flight requires that the flap to wing interface loads be sensed and monitored in real time to ensure that the wing structural load limits are not exceeded. This paper discusses the strain gage load calibration testing and load equation derivation methodology for the ACTE interface fittings. Both the left and right wing flap interfaces will be monitored and each contains four uniquely designed and instrumented flap interface fittings. The interface hardware design and instrumentation layout are discussed. Twenty one applied test load cases were developed using the predicted in-flight loads for the ACTE experiment.

Aeroelastic Response of the Adaptive Compliant Trailing Edge Transtition Section

NASA Technical Reports Server (NTRS)

Herrera, Claudia Y.; Spivey, Natalie D.; Lung, Shun-fat

2016-01-01

The Adaptive Compliant Trailing Edge demonstrator was a joint task under the Environmentally Responsible Aviation Project in partnership with the Air Force Research Laboratory and FlexSys, Inc. (Ann Arbor, Michigan), chartered by the National Aeronautics and Space Administration to develop advanced technologies that enable environmentally friendly aircraft, such as continuous mold-line technologies. The Adaptive Compliant Trailing Edge demonstrator encompassed replacing the Fowler flaps on the SubsoniC Aircraft Testbed, a Gulfstream III (Gulfstream Aerospace, Savannah, Georgia) aircraft, with control surfaces developed by FlexSys, Inc., a pair of uniquely-designed, unconventional flaps to be used as lifting surfaces during flight-testing to substantiate their structural effectiveness. The unconventional flaps consisted of a main flap section and two transition sections, inboard and outboard, which demonstrated the continuous mold-line technology. Unique characteristics of the transition sections provided a challenge to the airworthiness assessment for this part of the structure. A series of build-up tests and analyses were conducted to ensure the data required to support the airworthiness assessment were acquired and applied accurately. The transition sections were analyzed both as individual components and as part of the flight-test article assembly. Instrumentation was installed in the transition sections based on the analysis to best capture the in-flight aeroelastic response. Flight-testing was conducted and flight data were acquired to validate the analyses. This paper documents the details of the aeroelastic assessment and in-flight response of the transition sections of the unconventional Adaptive Compliant Trailing Edge flaps.

Large-Scale Wind-Tunnel Tests and Evaluation of the Low-Speed Performance of a 35 deg Sweptback Wing Jet Transport Model Equipped with a Blowing Boundary-Layer-Control Flap and Leading-Edge Slat

NASA Technical Reports Server (NTRS)

Hickey, David H.; Aoyagi, Kiyoshi

1960-01-01

A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.

Airframe noise

NASA Astrophysics Data System (ADS)

Crighton, David G.

1991-08-01

Current understanding of airframe noise was reviewed as represented by experiment at model and full scale, by theoretical modeling, and by empirical correlation models. The principal component sources are associated with the trailing edges of wing and tail, deflected trailing edge flaps, flap side edges, leading edge flaps or slats, undercarriage gear elements, gear wheel wells, fuselage and wing boundary layers, and panel vibration, together with many minor protrusions like radio antennas and air conditioning intakes which may contribute significantly to perceived noise. There are also possibilities for interactions between the various mechanisms. With current engine technology, the principal airframe noise mechanisms dominate only at low frequencies, typically less than 1 kHz and often much lower, but further reduction of turbomachinery noise in particular may make airframe noise the principal element of approach noise at frequencies in the sensitive range.

Blowing-Type Boundary-Layer Control as Applied to the Trailing-Edge Flaps of a 35 Degree Swept-Wing Airplane

NASA Technical Reports Server (NTRS)

Kelly, Mark W; Anderson, Seth B; Innis, Robert C

1958-01-01

A wind-tunnel investigation was made to determine the effects on the aerodynamic characteristics of a 35 degree swept-wing airplane of applying blowing-type boundary-layer control to the trailing-edge flaps. Flight tests of a similar airplane were then conducted to determine the effects of boundary-layer control on the handling qualities and operation of the airplane, particularly during landing and take-off. The wind-tunnel and flight tests indicated that blowing over the flaps produced large increases in flap lift increment, and significant increases in maximum lift. The use of blowing permitted reductions in the landing approach speeds of as much as 12 knots.

Flight control system development and flight test experience with the F-111 mission adaptive wing aircraft

NASA Technical Reports Server (NTRS)

Larson, R. R.

1986-01-01

The wing on the NASA F-111 transonic aircraft technology airplane was modified to provide flexible leading and trailing edge flaps. This wing is known as the mission adaptive wing (MAW) because aerodynamic efficiency can be maintained at all speeds. Unlike a conventional wing, the MAW has no spoilers, external flap hinges, or fairings to break the smooth contour. The leading edge flaps and three-segment trailing edge flaps are controlled by a redundant fly-by-wire control system that features a dual digital primary system architecture providing roll and symmetric commands to the MAW control surfaces. A segregated analog backup system is provided in the event of a primary system failure. This paper discusses the design, development, testing, qualification, and flight test experience of the MAW primary and backup flight control systems.

HSCT Ref-H Transonic Flap Data Base: Wind-Tunnel Test and Comparison with Theory

NASA Technical Reports Server (NTRS)

Vijgen, Paul M.

1999-01-01

In cooperation with personnel from the Boeing ANP Laboratory and NASA Langley, a performance test was conducted using the Reference-H 1.675% model ("NASA Modular Model") without nacelles at the NASA Langley 16-Ft Transonic Tunnel. The main objective of the test was to determine the drag reduction achievable with leading-edge and trailing-edge flaps deflected along the outboard wing span at transonic Mach numbers (M = 0.9 to 1.2) for purpose of preliminary design and for comparison with computational predictions. The obtained drag data with flap deflections for Mach numbers of 1.07 to 1.20 are unique for the Reference H wing. Four leading-edge and two trailing-edge flap deflection angles were tested at a mean-wing chord-Reynolds number of about 5.7 million. An outboard-wing leading-edge flap deflection of 81 provides a 4.5 percent drag reduction at M = 1.2 A = 0.2), and much larger values at lower Mach numbers with larger flap deflections. The present results for the baseline (no flaps deflected) compare reasonably well with previous Boeing and NASA Ref-H tunnel tests, including high-Reynolds number NTF results. Viscous CFD simulations using the OVERFLOW thin-layer N.S. method properly predict the observed trend in drag reduction at M = 1.2 as function of leading-edge flap deflection. Modified linear theory properly predicts the flap effects on drag at subsonic conditions (Aero2S code), and properly predicts the absolute drag for the 40 and 80 leading-edge deflection at M = 1.2 (A389 code).

Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One with Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One with a Flared Afterbody and All-Movable Controls

NASA Technical Reports Server (NTRS)

Robinson, Ross B; Bernot, Peter T

1958-01-01

An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 deg the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 deg flared afterbody. The tests were made through an angle-of-attack range of -2 deg to 20 deg at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge -flap configuration. The flared -after body configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift drag ratio.

Aerodynamic Characteristics at a Mach Number of 6.8 of Two Hypersonic Missile Configurations, One with Low-Aspect-Ratio Cruciform Fins and Trailing-Edge Flaps and One with a Flared Afterbody and All-Movable Controls

NASA Technical Reports Server (NTRS)

Bernot, P. T.; Robinson, R. B.

1958-01-01

An investigation has been made to determine the aerodynamic characteristics in pitch at a Mach number of 6.8 of hypersonic missile configurations with cruciform trailing-edge flaps and with all-movable control surfaces. The flaps were tested on a configuration having low-aspect-ratio cruciform fins with an apex angle of 5 degrees; the all-movable controls were mounted at the 46.7-percent body station on a configuration having a 10 degrees flared afterbody. The tests were made through an angle-of-attack range of -2 degrees to 20 degrees at zero sideslip in the Langley 11-inch hypersonic tunnel. The results indicated that the all-movable controls on the flared-afterbody model should be capable of producing much larger values of trim lift and of normal acceleration than the trailing-edge-flap configuration. The flared-afterbody configuration had considerably higher drag than the cruciform-fin model but only slightly lower values of lift-drag ratio.

Low-speed aerodynamic performance of an aspect-ratio-10 supercritical-wing transport model equipped with a full-span slat and part-span and full-span double-slotted flaps

NASA Technical Reports Server (NTRS)

Morgan, H. L., Jr.

1981-01-01

An investigation was conducted in the Langley 4 by 7 Meter Tunnel to determine the static longitudinal and lateral directional aerodynamic characteristics of an advanced aspect ratio 10 supercritical wing transport model equipped with a full span leading edge slat as well as part span and full span trailing edge flaps. This wide body transport model was also equipped with spoiler and aileron roll control surfaces, flow through nacelles, landing gear, and movable horizontal tails. Six basic wing configurations were tested: (1) cruise (slats and flaps nested), (2) climb (slats deflected and flaps nested), (3) part span flap, (4) full span flap, (5) full span flap with low speed ailerons, and (6) full span flap with high speed ailerons. Each of the four flapped wing configurations was tested with leading edge slat and trailing edge flaps deflected to settings representative of both take off and landing conditions. Tests were conducted at free stream conditions corresponding to Reynolds number of 0.97 to 1.63 x 10 to the 6th power and corresponding Mach numbers of 0.12 to 0.20, through an angle of attack range of 4 to 24, and a sideslip angle range of -10 deg to 5 deg. The part and full span wing configurations were also tested in ground proximity.

Aerodynamic effects of flexibility in flapping wings.

PubMed

Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P

2010-03-06

Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re approximately 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic insects and, to a limited extent, in understanding the aerodynamics of flapping insect wings.

Aerodynamic effects of flexibility in flapping wings

PubMed Central

Zhao, Liang; Huang, Qingfeng; Deng, Xinyan; Sane, Sanjay P.

2010-01-01

Recent work on the aerodynamics of flapping flight reveals fundamental differences in the mechanisms of aerodynamic force generation between fixed and flapping wings. When fixed wings translate at high angles of attack, they periodically generate and shed leading and trailing edge vortices as reflected in their fluctuating aerodynamic force traces and associated flow visualization. In contrast, wings flapping at high angles of attack generate stable leading edge vorticity, which persists throughout the duration of the stroke and enhances mean aerodynamic forces. Here, we show that aerodynamic forces can be controlled by altering the trailing edge flexibility of a flapping wing. We used a dynamically scaled mechanical model of flapping flight (Re ≈ 2000) to measure the aerodynamic forces on flapping wings of variable flexural stiffness (EI). For low to medium angles of attack, as flexibility of the wing increases, its ability to generate aerodynamic forces decreases monotonically but its lift-to-drag ratios remain approximately constant. The instantaneous force traces reveal no major differences in the underlying modes of force generation for flexible and rigid wings, but the magnitude of force, the angle of net force vector and centre of pressure all vary systematically with wing flexibility. Even a rudimentary framework of wing veins is sufficient to restore the ability of flexible wings to generate forces at near-rigid values. Thus, the magnitude of force generation can be controlled by modulating the trailing edge flexibility and thereby controlling the magnitude of the leading edge vorticity. To characterize this, we have generated a detailed database of aerodynamic forces as a function of several variables including material properties, kinematics, aerodynamic forces and centre of pressure, which can also be used to help validate computational models of aeroelastic flapping wings. These experiments will also be useful for wing design for small robotic insects and, to a limited extent, in understanding the aerodynamics of flapping insect wings. PMID:19692394

Low-speed investigation of effects of wing leading- and trailing-edge flap deflections and canard incidence on a fighter configuration equipped with a forward-swept wing

NASA Technical Reports Server (NTRS)

Gainer, T. G.; Mann, M. J.; Huffman, J. K.

1984-01-01

An advanced fighter configuration with a forward-swept wing of aspect ratio 3.28 is tested in the Langley 7 by 10 Foot High Speed Tunnel at a Mach number of 0.3. The wing has 29.5 degrees of forward sweep of the quarter chord line and is equipped with 15 percent chord leading edge and 30 percent chord trailing edge flaps. The canard is sweptback 45 degrees. Tests were made through a range of angle of attack from about -2 degrees to 22 degrees. Deflecting the flaps significantly improves the lift drag characteristics at the higher angles of attack. The canard is able to trim the configurations with different flap deflections over most of the range of angle of attack. The penalty in maximum lift coefficient due to trimming is about 0.10.

Longitudinal Aerodynamic Modeling of the Adaptive Compliant Trailing Edge Flaps on a GIII Airplane and Comparisons to Flight Data

NASA Technical Reports Server (NTRS)

Smith, Mark S.; Bui, Trong T.; Garcia, Christian A.; Cumming, Stephen B.

2016-01-01

A pair of compliant trailing edge flaps was flown on a modified GIII airplane. Prior to flight test, multiple analysis tools of various levels of complexity were used to predict the aerodynamic effects of the flaps. Vortex lattice, full potential flow, and full Navier-Stokes aerodynamic analysis software programs were used for prediction, in addition to another program that used empirical data. After the flight-test series, lift and pitching moment coefficient increments due to the flaps were estimated from flight data and compared to the results of the predictive tools. The predicted lift increments matched flight data well for all predictive tools for small flap deflections. All tools over-predicted lift increments for large flap deflections. The potential flow and Navier-Stokes programs predicted pitching moment coefficient increments better than the other tools.

Computational Investigations on the Effects of Gurney Flap on Airfoil Aerodynamics.

PubMed

Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram

2015-01-01

The present study comprises steady state, two-dimensional computational investigations performed on NACA 0012 airfoil to analyze the effect of Gurney flap (GF) on airfoil aerodynamics using k-ε RNG turbulence model of FLUENT. Airfoil with GF is analyzed for six different heights from 0.5% to 4% of the chord length, seven positions from 0% to 20% of the chord length from the trailing edge, and seven mounting angles from 30° to 120° with the chord. Computed values of lift and drag coefficients with angle of attack are compared with experimental values and good agreement is found at low angles of attack. In addition static pressure distribution on the airfoil surface and pathlines and turbulence intensities near the trailing edge are present. From the computational investigation, it is recommended that Gurney flaps with a height of 1.5% chord be installed perpendicular to chord and as close to the trailing edge as possible to obtain maximum lift enhancement with minimum drag penalty.

Computational Investigations on the Effects of Gurney Flap on Airfoil Aerodynamics

PubMed Central

Jain, Shubham; Sitaram, Nekkanti; Krishnaswamy, Sriram

2015-01-01

The present study comprises steady state, two-dimensional computational investigations performed on NACA 0012 airfoil to analyze the effect of Gurney flap (GF) on airfoil aerodynamics using k-ε RNG turbulence model of FLUENT. Airfoil with GF is analyzed for six different heights from 0.5% to 4% of the chord length, seven positions from 0% to 20% of the chord length from the trailing edge, and seven mounting angles from 30° to 120° with the chord. Computed values of lift and drag coefficients with angle of attack are compared with experimental values and good agreement is found at low angles of attack. In addition static pressure distribution on the airfoil surface and pathlines and turbulence intensities near the trailing edge are present. From the computational investigation, it is recommended that Gurney flaps with a height of 1.5% chord be installed perpendicular to chord and as close to the trailing edge as possible to obtain maximum lift enhancement with minimum drag penalty. PMID:27347517

Predicted Aerodynamic Characteristics of a NACA 0015 Airfoil Having a 25% Integral-Type Trailing Edge Flap

NASA Technical Reports Server (NTRS)

Hassan, Ahmed

1999-01-01

Using the two-dimensional ARC2D Navier-Stokes flow solver analyses were conducted to predict the sectional aerodynamic characteristics of the flapped NACA-0015 airfoil section. To facilitate the analyses and the generation of the computational grids, the airfoil with the deflected trailing edge flap was treated as a single element airfoil with no allowance for a gap between the flap's leading edge and the base of the forward portion of the airfoil. Generation of the O-type computational grids was accomplished using the HYGRID hyperbolic grid generation program. Results were obtained for a wide range of Mach numbers, angles of attack and flap deflections. The predicted sectional lift, drag and pitching moment values for the airfoil were then cast in tabular format (C81) to be used in lifting-line helicopter rotor aerodynamic performance calculations. Similar were also generated for the flap. Mathematical expressions providing the variation of the sectional lift and pitching moment coefficients for the airfoil and for the flap as a function of flap chord length and flap deflection angle were derived within the context of thin airfoil theory. The airfoil's sectional drag coefficient were derived using the ARC2D drag predictions for equivalent two dimensional flow conditions.

Implementation of a Trailing-Edge Flap Analysis Model in the NASA Langley CAMRAD.MOD1/Hires Program

NASA Technical Reports Server (NTRS)

Charles, Bruce

1999-01-01

Continual advances in rotorcraft performance, vibration and acoustic characteristics are being sought by rotary-wing vehicle manufacturers to improve efficiency, handling qualities and community noise acceptance of their products. The rotor system aerodynamic and dynamic behavior are among the key factors which must be addressed to meet the desired goals. Rotor aerodynamicists study how airload redistribution impacts performance and noise, and seek ways to achieve better airload distribution through changes in local aerodynamic response characteristics. One method currently receiving attention is the use of trailing-edge flaps mounted on the rotor blades to provide direct control of a portion of the spanwise lift characteristics. The following work describes the incorporation of a trailing-edge flap model in the CAMRAD.Mod1/FHUS comprehensive rotorcraft analysis code. The CAM-RAD.Mod1/HIRES analysis consists of three separate executable codes. These include the comprehensive trim analysis, CAMRAD.Mod1, the Indicial Post-Processor, IPP, for high resolution airloads, and AIRFOIL, which produces the rotor airfoil tables from input airfoil section characteristics. The modifications made to these components permitting analysis of flapped rotor configurations are documented herein along with user instructions detailing the new input variables and operational notes.

Pressure Distribution Over a Symmetrical Airfoil Section with Trailing Edge Flap

NASA Technical Reports Server (NTRS)

Jacobs, Eastman N; Pinkerton, Robert M

1931-01-01

Measurements were made to determine the distribution of pressure over one section of an R. A. F. 30 (symmetrical) airfoil with trailing edge flaps. In order to study the effect of scale measurements were made with air densities of approximately 1 and 20 atmospheres. Isometric diagrams of pressure distribution are given to show the effect of change in incidence, flap displacement, and scale upon the distribution. Plots of normal force coefficient versus angle of attack for different flap displacements are given to show the effect of a displaced flap. Plots are given of both the experimental and theoretical characteristic coefficients versus flap angle, in order to provide a comparison with the theory. It is concluded that for small flap displacements the agreement for the pitching and hinge moments is such that it warrants the use of the theoretical parameters. However, the agreement for the lift is not as good, particularly for the smaller flaps. In an appendix, an example is given of the calculation of the load and moments on an airfoil with hinged flap from these parameters.

Extreme load alleviation using industrial implementation of active trailing edge flaps in a full design load basis

NASA Astrophysics Data System (ADS)

Barlas, Thanasis; Pettas, Vasilis; Gertz, Drew; Madsen, Helge A.

2016-09-01

The application of active trailing edge flaps in an industrial oriented implementation is evaluated in terms of capability of alleviating design extreme loads. A flap system with basic control functionality is implemented and tested in a realistic full Design Load Basis (DLB) for the DTU 10MW Reference Wind Turbine (RWT) model and for an upscaled rotor version in DTU's aeroelastic code HAWC2. The flap system implementation shows considerable potential in reducing extreme loads in components of interest including the blades, main bearing and tower top, with no influence on fatigue loads and power performance. In addition, an individual flap controller for fatigue load reduction in above rated power conditions is also implemented and integrated in the general controller architecture. The system is shown to be a technology enabler for rotor upscaling, by combining extreme and fatigue load reduction.

Pitch, roll, and yaw moment generator for insect-like tailless flapping-wing MAV

NASA Astrophysics Data System (ADS)

Phan, Hoang Vu; Park, Hoon Cheol

2016-04-01

In this work, we proposed a control moment generator, which is called Trailing Edge Change (TEC) mechanism, for attitudes change in hovering insect-like tailless flapping-wing MAV. The control moment generator was installed to the flapping-wing mechanism to manipulate the wing kinematics by adjusting the wing roots location symmetrically or asymmetrically. As a result, the mean aerodynamic force center of each wing is relocated and control moments are generated. The three-dimensional wing kinematics captured by three synchronized high-speed cameras showed that the flapping-wing MAV can properly modify the wing kinematics. In addition, a series of experiments were performed using a multi-axis load cell to evaluate the forces and moments generation. The measurement demonstrated that the TEC mechanism produced reasonable amounts of pitch, roll and yaw moments by shifting position of the trailing edges at the wing roots of the flapping-wing MAV.

High-Lift Systems on Commercial Subsonic Airliners

NASA Technical Reports Server (NTRS)

Rudolph, Peter K. C.

1996-01-01

The early breed of slow commercial airliners did not require high-lift systems because their wing loadings were low and their speed ratios between cruise and low speed (takeoff and landing) were about 2:1. However, even in those days the benefit of high-lift devices was recognized. Simple trailing-edge flaps were in use, not so much to reduce landing speeds, but to provide better glide-slope control without sideslipping the airplane and to improve pilot vision over the nose by reducing attitude during low-speed flight. As commercial-airplane cruise speeds increased with the development of more powerful engines, wing loadings increased and a real need for high-lift devices emerged to keep takeoff and landing speeds within reasonable limits. The high-lift devices of that era were generally trailing-edge flaps. When jet engines matured sufficiently in military service and were introduced commercially, airplane speed capability had to be increased to best take advantage of jet engine characteristics. This speed increase was accomplished by introducing the wing sweep and by further increasing wing loading. Whereas increased wing loading called for higher lift coefficients at low speeds, wing sweep actually decreased wing lift at low speeds. Takeoff and landing speeds increased on early jet airplanes, and, as a consequence, runways worldwide had to be lengthened. There are economical limits to the length of runways; there are safety limits to takeoff and landing speeds; and there are speed limits for tires. So, in order to hold takeoff and landing speeds within reasonable limits, more powerful high-lift devices were required. Wing trailing-edge devices evolved from plain flaps to Fowler flaps with single, double, and even triple slots. Wing leading edges evolved from fixed leading edges to a simple Krueger flap, and from fixed, slotted leading edges to two- and three-position slats and variable-camber (VC) Krueger flaps. The complexity of high-lift systems probably peaked on the Boeing 747, which has a VC Krueger flap and triple-slotted, inboard and outboard trailing-edge flaps. Since then, the tendency in high-lift system development has been to achieve high levels of lift with simpler devices in order to reduce fleet acquisition and maintenance costs. The intent of this paper is to: (1) review available high-lift devices, their functions, and design criteria; (2) appraise high-lift systems presently in service on commercial air liners; (3) present personal study results on high-lift systems; (4) develop a weight and cost model for high-lift systems; and (5) discuss the development tendencies of future high-lift systems.

Theoretical characteristics of two-dimensional supersonic control surfaces

NASA Technical Reports Server (NTRS)

Morrissette, Robert R; Oborny, Lester F

1951-01-01

The "Busemann second-order-approximation theory" for the pressure distribution over a two-dimensional airfoil in supersonic flow was used to determine some of the aerodynamic characteristics of uncambered symmetrical parabolic and double-wedge airfoils with leading-edge and trailing-edge flaps. The characteristics presented and discussed in this paper are: flap effectiveness factor, rate of change of hinge-moment coefficient with flap deflection, rate of change of the pitching-moment coefficient with flap deflection, rate of change of the pitching-moment coefficient about the mid chord with flap deflection, and the location of the center of pressure of the airfoil-flap combination.

Influencing the aerodynamics of the ACFA2020 aircraft with flap and trailing edge device oscillations

NASA Astrophysics Data System (ADS)

Meyer, M.; Breitsamter, Ch.

2013-12-01

The influence of an oscillating aileron and trailing edge device on the unsteady aerodynamics of a blended wing body (BWB) aircraft configuration with high-fidelity time-accurate Euler simulations has been investigated. Steady results show an unequally-distributed lift distribution in spanwise direction with a particularly severe shock at cruise conditions on the outboard wing. Unsteady oscillations of the outboardlocated aileron are able to influence the local and global aerodynamics. The oscillation of the trailing edge device designed to be at trailing edge of the aileron does not show any great effect on neither local nor global aerodynamics.

Active Control of Separation From the Flap of a Supercritical Airfoil

NASA Technical Reports Server (NTRS)

Melton, LaTunia Pack; Yao, Chung-Sheng; Seifert, Avi

2006-01-01

Zero-mass-flux periodic excitation was applied at several regions on a simplified high-lift system to delay the occurrence of flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge flap and a 25% chord simply hinged trailing edge flap. Detailed flow features were measured in an attempt to identify optimal actuator placement. The measurements included steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization, and particle image velocimetry (PIV). The current paper describes the application of active separation control at several locations on the deflected trailing edge flap. High frequency (F(+) approximately equal to 10) and low frequency amplitude modulation (F(+) sub AM approximately equal to 1) of the high frequency excitation were used for control. It was noted that the same performance gains were obtained with amplitude modulation and required only 30% of the momentum input required by pure sine excitation.

Variable camber rotor study

NASA Technical Reports Server (NTRS)

Dadone, L.; Cowan, J.; Mchugh, F. J.

1982-01-01

Deployment of variable camber concepts on helicopter rotors was analytically assessed. It was determined that variable camber extended the operating range of helicopters provided that the correct compromise can be obtained between performance/loads gains and mechanical complexity. A number of variable camber concepts were reviewed on a two dimensional basis to determine the usefulness of leading edge, trailing edge and overall camber variation schemes. The most powerful method to vary camber was through the trailing edge flaps undergoing relatively small motions (-5 deg to +15 deg). The aerodynamic characteristics of the NASA/Ames A-1 airfoil with 35% and 50% plain trailing edge flaps were determined by means of current subcritical and transonic airfoil design methods and used by rotor performance and loads analysis codes. The most promising variable camber schedule reviewed was a configuration with a 35% plain flap deployment in an on/off mode near the tip of a blade. Preliminary results show approximately 11% reduction in power is possible at 192 knots and a rotor thrust coefficient of 0.09. The potential demonstrated indicates a significant potential for expanding the operating envelope of the helicopter. Further investigation into improving the power saving and defining the improvement in the operational envelope of the helicopter is recommended.

Multi-Element Airfoil System

NASA Technical Reports Server (NTRS)

Turner, Travis L. (Inventor); Khorrami, Mehdi R. (Inventor); Lockard, David P. (Inventor); McKenney, Martin J. (Inventor); Atherley, Raymond D. (Inventor); Kidd, Reggie T. (Inventor)

2014-01-01

A multi-element airfoil system includes an airfoil element having a leading edge region and a skin element coupled to the airfoil element. A slat deployment system is coupled to the slat and the skin element, and is capable of deploying and retracting the slat and the skin element. The skin element substantially fills the lateral gap formed between the slat and the airfoil element when the slat is deployed. The system further includes an uncoupling device and a sensor to remove the skin element from the gap based on a critical angle-of-attack of the airfoil element. The system can alternatively comprise a trailing edge flap, where a skin element substantially fills the lateral gap between the flap and the trailing edge region of the airfoil element. In each case, the skin element fills a gap between the airfoil element and the deployed flap or slat to reduce airframe noise.

The effect of chordwise flexibility on flapping foil propulsion in quiescent fluid

NASA Astrophysics Data System (ADS)

Shinde, Sachin; Arakeri, Jaywant

2010-11-01

Motivated to understand the role of wing flexibility of flying creatures during hovering, we experimentally study the effect of chordwise flexibility on the flow generated in quiescent fluid by a sinusoidally pitching rigid symmetrical foil with a flexible flap attached at the trailing edge. This foil produces a narrow, coherent jet containing reverse Karman vortex street, and a corresponding thrust. The thrust and flow is similar to that produced by a hovering bird or insect, however the mechanism seems to be different from known hovering mechanisms. Novelty of the present hovering mechanism is that the thrust generation is due to the coordinated pushing action of rigid foil and flexible flap. We identify the flow and vortex generation mechanism. This foil produces jet flows over a range of flapping frequencies and amplitudes. In contrast, the foil without flap i.e. with rigid trailing edge produces a weak, divergent jet that meanders randomly. Appending a flexible flap to the foil suppresses jet-meandering and strengthens the jet. Flexibility of flap is crucial in determining the flow structure. This study is useful in designing MAVs and thrusters.

Lifting-surface theory for calculating the loading induced on a wing by a flap

NASA Technical Reports Server (NTRS)

Johnson, W. A.

1972-01-01

A method is described for using lifting-surface theory to obtain the pressure distribution on a wing with a trailing-edge flap or control surface. The loading has a logarithmic singularity at the flap edges, which may be determined directly by the method of matched asymptotic expansions. Expressions are given for the singular flap loading for various flap hinge line and side edge geometries, both for steady and unsteady flap deflection. The regular part of the flap loading must be obtained by inverting the lifting-surface-theory integral equation relating the pressure and the downwash on the wing: procedures are described to accomplish this for a general wing and flap geometry. The method is applied to several example wings, and the results are compared with experimental data. Theory and test correlate well.

The Effect of Split Trailing-edge Wing Flaps on the Aerodynamic Characteristics of a Parasol Monoplane

NASA Technical Reports Server (NTRS)

Wallace, Rudolf, N

1933-01-01

This paper presents the results of tests conducted in the N.A.C.A. full-scale wind tunnel on a Fairchild F-22 airplane equipped with a special wing having split trailing-edge flaps. The flaps extended over the outer 90 percent of the wing span, and were of the fixed-hinge type having a width equal to 20 percent of the wing chord. The results show that with a flap setting of 59 degrees the maximum lift of the wing was increased 42 percent, and that the flaps increased the range of available gliding angles from 2.7 degrees to 7.0 degrees. Deflection of the split flaps did not increase the stalling angle or seriously affect the longitudinal balance of the airplane. With flaps down the landing speed of the airplane is decreased, but the calculated climb and level-flight performance is inferior to that with the normal wing. Calculations indicate that the take-off distance required to clear an obstacle 100 feet high is not affected by flap settings from 0 degrees to 20 degrees but is greatly increased by larger flap angles.

Approach for Structurally Clearing an Adaptive Compliant Trailing Edge Flap for Flight

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Lokos, William A.; Cruz, Josue; Crampton, Glen; Stephens, Craig A.; Kota, Sridhar; Ervin, Gregory; Flick, Pete

2015-01-01

The Adaptive Compliant Trailing Edge (ACTE) flap was flown on the National Aeronautics and Space Administration (NASA) Gulfstream GIII testbed at the NASA Armstrong Flight Research Center. This smoothly curving flap replaced the existing Fowler flaps creating a seamless control surface. This compliant structure, developed by FlexSys Inc. in partnership with the Air Force Research Laboratory, supported NASA objectives for airframe structural noise reduction, aerodynamic efficiency, and wing weight reduction through gust load alleviation. A thorough structures airworthiness approach was developed to move this project safely to flight. A combination of industry and NASA standard practice require various structural analyses, ground testing, and health monitoring techniques for showing an airworthy structure. This paper provides an overview of compliant structures design, the structural ground testing leading up to flight, and the flight envelope expansion and monitoring strategy. Flight data will be presented, and lessons learned along the way will be highlighted.

The calculated effect of trailing-edge flaps on the take-off of flying boats

NASA Technical Reports Server (NTRS)

Parkinson, J E; Bell, J W

1934-01-01

The results of take-off calculations are given for an application of simple trailing-edge flaps to two hypothetical flying boats, one having medium wing and power loading and consequently considerable excess of thrust over total resistance during the take-off run, the other having high wing and power loading and a very low excess thrust. For these seaplanes the effect of downward flap settings was: (1) to increase the total resistance below the stalling speed, (2) to decrease the get-away speed, (3) to improve the take-off performance of the seaplane having considerable excess thrust, and (4) to hinder the take-off of the seaplane having low excess thrust. It is indicated that flaps would allow a decrease in the high angles of wing setting necessary with most seaplanes, provided that the excess thrust is not too low.

A wind tunnel investigation of the wake near the trailing edge of a deflected externally blown flap. [on a jet powered STOL transport aircraft

NASA Technical Reports Server (NTRS)

Johnson, W. G., Jr.; Kardas, G. E.

1974-01-01

The model tested was a general research model of a swept-wing, jet-powered STOL transport with externally blown flaps. The model was tested with four engine simulators mounted on pylons under the wing. Tests were conducted in the V/STOL tunnel over an angle of attack range of 0 deg to 16 deg and a thrust coefficient range from 0 to approximately 4 at a Reynolds number of 0.461 x 1 million based on the wing reference chord. The results of this investigation are presented primarily as plots of the individual velocity vectors obtained from the wake survey. These data are used to extend an earlier analysis to isolate the effects of the engine thrust on the behavior of the flow at the flap trailing edge. Results of a comparison with a jet-flap theory are also shown.

Vortex Flap Technology: a Stability and Control Assessment

NASA Technical Reports Server (NTRS)

Carey, K. M.; Erickson, G. E.

1984-01-01

A comprehensive low-speed wind tunnel investigation was performed of leading edge vortex flaps applied to representative aircraft configurations. A determination was made of the effects of analytically- and empirically-designed vortex flaps on the static longitudinal and lateral-directional aerodynamics, stability, and control characteristics of fighter wings having leading-edge sweep angles of 45 to 76.5 degrees. The sensitivity to several configuration modifications was assessed, which included the effects of flap planform, leading- and trailing-edge flap deflection angles, wing location on the fuselage, forebody strakes, canards, and centerline and outboard vertical tails. Six-component forces and moments, wing surface static pressure distributions, and surface flow patterns were obtained using the Northrop 21- by 30-inch low-speed wind tunnel.

Investigation of trailing-edge-flap, spanwise-blowing concepts on an advanced fighter configuration

NASA Technical Reports Server (NTRS)

Paulson, J. W., Jr.; Quinto, P. F.; Banks, D. W.

1984-01-01

The aerodynamic effects of spanwise blowing on the trailing edge flap of an advanced fighter aircraft configuration were determined in the 4 by 7 Meter Tunnel. A series of tests were conducted with variations in spanwise-blowing vector angle, nozzle exit area, nozzle location, thrust coefficient, and flap deflection in order to determine a superior configuration for both an underwing cascade concept and an overwing port concept. This screening phase of the testing was conducted at a nominal approach angle of attack from 12 deg to 16 deg; and then the superior configurations were tested over a more complete angle of attack range from 0 deg to 20 deg at tunnel free stream dynamic pressures from 20 to 40 lbf/sq ft at thrust coefficients from 0 to 2.

AMELIA CESTOL Test: Acoustic Characteristics of Circulation Control Wing with Leading-and Trailing-Edge Slot Blowing

NASA Technical Reports Server (NTRS)

Horne, Clifton; Burnside, Nathan J.

2013-01-01

Aeroacoustic measurements of the 11 % scale full-span AMELIA CESTOL model with leading- and trailing-edge slot blowing circulation control (CCW) wing were obtained during a recent test in the Arnold Engineering Development Center 40- by 80-Ft. Wind Tunnel at NASA Ames Research Center, Sound levels and spectra were acquired with seven in-flow microphones and a 48-element phased microphone array for a variety of vehicle configurations, CCW slot flow rates, and forward speeds, Corrections to the measurements and processing are in progress, however the data from selected configurations presented in this report confirm good measurement quality and dynamic range over the test conditions, Array beamform maps at 40 kts tunnel speed show that the trailing edge flap source is dominant for most frequencies at flap angles of 0deg and 60deg, The overall sound level for the 60deg flap was similar to the 0deg flap for most slot blowing rates forward of 90deg incidence, but was louder by up to 6 dB for downstream angles, At 100 kts, the in-flow microphone levels were louder than the sensor self-noise for the higher blowing rates, while passive and active background noise suppression methods for the microphone array revealed source levels as much as 20 dB lower than observed with the in-flow microphones,

Flaperon Modification Effect on Jet-Flap Interaction Noise Reduction for Chevron Nozzles

NASA Technical Reports Server (NTRS)

Thomas, Russell H.; Mengle, Vinod G.; Stoker, Robert W.; Brusniak, Leon; Elkoby, Ronen

2007-01-01

Jet-flap interaction (JFI) noise can become an important component of far field noise when a flap is immersed in the engine propulsive stream or is in its entrained region, as in approach conditions for under-the-wing engine configurations. We experimentally study the effect of modifying the flaperon, which is a high speed aileron between the inboard and outboard flaps, at both approach and take-off conditions using scaled models in a free jet. The flaperon modifications were of two types: sawtooth trailing edge and mini vortex generators (vg s). Parametric variations of these two concepts were tested with a round coaxial nozzle and an advanced chevron nozzle, with azimuthally varying fan chevrons, using both far field microphone arrays and phased microphone arrays for source diagnostics purposes. In general, the phased array results corroborated the far field results in the upstream quadrant pointing to JFI near the flaperon trailing edge as the origin of the far field noise changes. Specific sawtooth trailing edges in conjunction with the round nozzle gave marginal reduction in JFI noise at approach, and parallel co-rotating mini-vg s were somewhat more beneficial over a wider range of angles, but both concepts were noisier at take-off conditions. These two concepts had generally an adverse JFI effect when used in conjunction with the advanced chevron nozzle at both approach and take-off conditions.

Large-Scale Wind-Tunnel Tests of an Airplane Model with an Unswept, Tilt Wing of Aspect Ratio 5.5, and with Four Propellers and Blowing Flaps

NASA Technical Reports Server (NTRS)

Weiberg, James A.; Holzhauser, Curt A.

1961-01-01

Tests were made of a large-scale tilt-wing deflected-slipstream VTOL airplane with blowing-type BLC trailing-edge flaps. The model was tested with flap deflections of 0 deg. without BLC, 50 deg. with and without BLC, and 80 deg. with BLC for wing-tilt angles of 0, 30, and 50 deg. Included are results of tests of the model equipped with a leading-edge flap and the results of tests of the model in the presence of a ground plane.

Aerodynamic Characteristics of a 0.5-Scale Model of the Fairchild XSAM-N-2 Lark Missile at High Subsonic Speeds

NASA Technical Reports Server (NTRS)

Martin, Andrew; Hunter, Harlo A.

1949-01-01

An investigation was conducted to determine the longitudinal- and lateral-stability characteristics of a 0.5-scale moue1 of the Fairchild Lark missile, The model was tested with 0 deg and with 22.5 deg of roll. Three horizontal wings having NACA 16-009, 16-209, and 64A-209 sections were tested. Pressures were measured on both pointed and blunt noses. The wind-tunnel-test data indicate that rolling the missile 22.5 deg. had no serious effect on the static longitudinal stability. The desired maneuvering acceleration could not be attained with any of the horizontal wings tested, even with the horizontal wing flaps deflected 50 deg. The flaps on the 64A-209 wing (with small trailing-edge angles and flat sides) were effective at all flap deflections, while the flaps on the 16-series wings (with large trailing-edge angles) lost effectiveness at small flap deflections. The data showed that rolling moment existed when the vertical wing flaps were deflected with the model at other than zero angle of attack. A similar rolling moment probably would be found . with the horizontal wing flaps deflected and the model yawed.

A PIV Study of Baseline and Controlled Flow over the Highly Deflected Flap of a Generic Low Aspect Ratio Trapezoidal Wing

NASA Astrophysics Data System (ADS)

Tewes, Philipp; Genschow, Konstantin; Little, Jesse; Wygnanski, Israel

2017-11-01

A detailed flow survey using PIV was conducted over a highly-deflected flap (55°) of a low-aspect ratio trapezoidal wing. The wing section is a NACA 0012 with 45° sweep at both the leading and trailing edges, an aspect ratio of 1.5 and a taper ratio of 0.27. The main element is equipped with 7 equally spaced fluidic oscillators, covering the inner 60 % of the span, located near the flap hinge. Experiments were carried out at 0° and 8° incidence at a Reynolds number of 1.7 .106 for both baseline and active flow control (AFC) cases. Velocity ISO-surfaces, x-vorticity and streamlines are analyzed / discussed. A flap leading edge vortex governs the baseline flow field for 0°. This vortical structure interacts with the jets emitted by the actuators (Cμ = 1 %). Its development is hampered and the vortex is redirected toward the trailing edge resulting in a CL increase. At 8°, the dominant flap leading edge vortex could not be detected and is believed to have already merged with the tip vortex. AFC attached the flow over the flap and enhanced the lift by up to 20 % while maintaining longitudinal stability. The dominant flow features in the AFC cases are actuator-generated streamwise vortices which appear stronger at 8°. This work was supported by the Office of Naval Research under ONR Grant No. N00014-14-1-0387.

Improvement of maneuver aerodynamics by spanwise blowing

NASA Technical Reports Server (NTRS)

Erickson, G. E.; Campbell, J. F.

1977-01-01

Spanwise blowing was used to test a generalized wind-tunnel model to investigate component concepts in order to provide improved maneuver characteristics for advanced fighter aircraft. Primary emphasis was placed on performance, stability, and control at high angles of attack and subsonic speeds. Test data were obtained in the Langley high speed 7 by 10 foot tunnel at free stream Mach numbers up to 0.50 for a range of model angles of attack, jet momentum coefficients, and leading and trailing edge flap deflection angles. Spanwise blowing on a 44 deg swept trapezoidal wing resulted in leading edge vortex enhancement with subsequent large vortex induced lift increments and drag polar improvements at the higher angles of attack. Small deflections of a leading edge flap delayed these lift and drag benefits to higher angles of attack. In addition, blowing was more effective at higher Mach numbers. Spanwise blowing in conjunction with a deflected trailing edge flap resulted in lift and drag benefits that exceeded the summation of the effects of each high lift device acting alone. Asymmetric blowing was an effective lateral control device at the higher angles of attack.

Gurney flap—Lift enhancement, mechanisms and applications

NASA Astrophysics Data System (ADS)

Wang, J. J.; Li, Y. C.; Choi, K.-S.

2008-01-01

Since its invention by a race car driver Dan Gurney in 1960s, the Gurney flap has been used to enhance the aerodynamics performance of subsonic and supercritical airfoils, high-lift devices and delta wings. In order to take stock of recent research and development of Gurney flap, we have carried out a review of the characteristics and mechanisms of lift enhancement by the Gurney flap and its applications. Optimum design of the Gurney flap is also summarized in this paper. For the Gurney flap to be effective, it should be mounted at the trailing edge perpendicular to the chord line of airfoil or wing. The flap height must be of the order of local boundary layer thickness. For subsonic airfoils, an additional Gurney flap increases the pressure on the upstream surface of the Gurney flap, which increases the total pressure of the lower surface. At the same time, a long wake downstream of the flap containing a pair of counter-rotating vortices can delay or eliminate the flow separation near the trailing edge on the upper surface. Correspondingly, the total suction on the airfoil is increased. For supercritical airfoils, the lift enhancement of the Gurney flap mainly comes from its ability to shift the shock on the upper surface in the downstream. Applications of the Gurney flap to modern aircraft design are also discussed in this review.

Flight Investigation of the Low-Speed Characteristics of a 45 deg Swept-Wing Fighter-Type Airplane with Blowing Boundary-Layer Control Applied to the Leading- and Trailing-Edge Flaps

NASA Technical Reports Server (NTRS)

Quigley, Hervey C.; Anderson, Seth B.; Innis, Robert C.

1960-01-01

A flight investigation has been conducted to study how pilots use the high lift available with blowing-type boundary-layer control applied to the leading- and trailing-edge flaps of a 45 deg. swept-wing airplane. The study includes documentation of the low-speed handling qualities as well as the pilots' evaluations of the landing-approach characteristics. All the pilots who flew the airplane considered it more comfortable to fly at low speeds than any other F-100 configuration they had flown. The major improvements noted were the reduced stall speed, the improved longitudinal stability at high lift, and the reduction in low-speed buffet. The study has shown the minimum comfortable landing-approach speeds are between 120.5 and 126.5 knots compared to 134 for the airplane with a slatted leading edge and the same trailing-edge flap. The limiting factors in the pilots' choices of landing-approach speeds were the limits of ability to control flight-path angle, lack of visibility, trim change with thrust, low static directional stability, and sluggish longitudinal control. Several of these factors were found to be associated with the high angles of attack, between 13 deg. and 15 deg., required for the low approach speeds. The angle of attack for maximum lift coefficient was 28 deg.

Airframe Noise Reduction Studies and Clean-Airframe Noise Investigation

NASA Technical Reports Server (NTRS)

Fink, M. R.; Bailey, D. A.

1980-01-01

Acoustic wind tunnel tests were conducted of a wing model with modified leading edge slat and trailing edge flap. The modifications were intended to reduce the surface pressure response to convected turbulence and thereby reduce the airframe noise without changing the lift at constant incidence. Tests were conducted at 70.7 and 100 m/sec airspeeds, with Reynolds numbers 1.5 x 10 to the 6th power and 2.1 x 10 to the 6th power. Considerable reduction of noise radiation from the side edges of a 40 deflection single slotted flap was achieved by modification to the side edge regions or the leading edge region of the flap panel. Total far field noise was reduced 2 to 3 dB over several octaves of frequency. When these panels were installed as the aft panel of a 40 deg deflection double slotted flap, 2 dB noise reduction was achieved.

Investigation of Vortex Flaps and Other Flow Control Devices on Generic High-Speed Civil Transport Planforms

NASA Technical Reports Server (NTRS)

Kjerstad, Kevin J.; Campbell, Bryan A.; Gile, Brenda E.; Kemmerly, Guy T.

1999-01-01

A parametric cranked delta planform study has been conducted in the Langley 14- by 22-Foot Subsonic Tunnel with the following objectives: (1) to evaluate the vortex flap design methodology for cranked delta wings, (2) to determine the influence of leading-edge sweep and the outboard wing on vortex flap effectiveness, (3) to evaluate novel flow control concepts, and (4) to validate unstructured grid Euler computer code predictions with modeled vortex and trailing-edge flaps. Two families of cranked delta planforms were investigated. One family had constant aspect ratio, while the other had a constant nondimensional semispan location of the leading-edge break. The inboard leading-edge sweep of the planforms was varied between 68 deg., 71 deg., and 74 deg., while outboard leading-edge sweep was varied between 48 deg. and 61 deg. Vortex flaps for the different planforms were designed by an analytical vortex flap design method. The results indicate that the effectiveness of the vortex flaps was only slightly influenced by the variations in the parametric planforms. The unstructured grid Euler computer code was successfully used to model the configurations with vortex flaps. The vortex trap concept was successfully demonstrated.

Experimental Test Results of the Energy Efficient Transport (EET) Flap-Edge Vortex Model in the Langley Low-Turbulence Pressure Tunnel

NASA Technical Reports Server (NTRS)

Morgan, Harry L., Jr.

2002-01-01

This report presents the results of a test conducted in the Langley Low-Turbulence Pressure Tunnel to measure the flow field properties of a flap-edge vortex. The model was the EET (Energy Efficient Transport) Flap-Edge Vortex Model, which consists of a main element and a part-span, single-slotted trailing-edge flap. The model surface was instrumented with several chordwise and spanwise rows of pressure taps on each element. The off-body flow field velocities were to be measured in several planes perpendicular to the flap edge with a laser velocimetry system capable of measuring all three components in coincidence. However, due to seeding difficulties, the preliminary laser data did not have sufficient accuracy to be suitable for presentation; therefore, this report presents only the tabulated and plotted surface pressure data. In addition, the report contains a detail description of the model which can be used to generate accurate CFD grid structures.

Evaluation of the Hinge Moment and Normal Force Aerodynamic Loads from a Seamless Adaptive Compliant Trailing Edge Flap in Flight

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Cruz, Josue; Lung, Shun-Fat; Kota, Sridhar; Ervin, Gregory; Lu, Kerr-Jia; Flick, Pete

2016-01-01

A seamless adaptive compliant trailing edge (ACTE) flap was demonstrated in flight on a Gulfstream III aircraft at the NASA Armstrong Flight Research Center. The trailing edge flap was deflected between minus 2 deg up and plus 30 deg down in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The safety-of-flight parameters for the ACTE flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. The attachment fittings connecting the flap to the aircraft wing rear spar were instrumented with strain gages and calibrated using known loads for measuring hinge moment and normal force loads in flight. The interface hardware instrumentation layout and load calibration are discussed. Twenty-one applied calibration test load cases were developed for each individual fitting. The 2-sigma residual errors for the hinge moment was calculated to be 2.4 percent, and for normal force was calculated to be 7.3 percent. The hinge moment and normal force generated by the ACTE flap with a hinge point located at 26-percent wing chord were measured during steady state and symmetric pitch maneuvers. The loads predicted from analysis were compared to the loads observed in flight. The hinge moment loads showed good agreement with the flight loads while the normal force loads calculated from analysis were over-predicted by approximately 20 percent. Normal force and hinge moment loads calculated from the pressure sensors located on the ACTE showed good agreement with the loads calculated from the installed strain gages.

Lift Augmentation on a Delta Wing via Leading Edge Fences and the Gurney Flap

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.; Tso, Jin

1993-01-01

Wind tunnel tests have been conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve by as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

Experimental study of the effect on span loading on aircraft wakes

NASA Technical Reports Server (NTRS)

Corsiglia, V. R.; Rossow, V. J.; Ciffone, D. L.

1975-01-01

Measurements were made in the NASA-Ames 40- by 80-foot wind tunnel of the rolling moment induced on a following model in the wake 13.6 spans behind a subsonic transport model for a variety of trailing edge flap settings of the generator. It was found that the rolling moment on the following model was reduced substantially, compared to the conventional landing configuration, by reshaping the span loading on the generating model to approximate a span loading, found in earlier studies, which resulted in reduced wake velocities. This was accomplished by retracting the outboard trailing edge flaps. It was concluded, based on flow visualization conducted in the wind tunnel as well as in a water tow facility, that this flap arrangement redistributes the vorticity shed by the wing along the span to form three vortex pairs that interact to disperse the wake.

Subsonic Investigation of Leading-Edge Flaps Designed for Vortex- and Attached-Flow on a High-Speed Civil Transport Configuration

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Kemmerly, Guy T.; Kjerstad, Kevin J.; Lessard, Victor R.

1999-01-01

A wind tunnel investigation of two separate leading-edge flaps, designed for vortex and attached-flow, respectively, were conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.12 to 0.27, with corresponding chord Reynolds numbers of 2.50 x 10 (sup 6) to 5.50 x 10 (sup 6). Variations of the leading-edge flap deflection angle were tested with outboard leading-edge flaps deflected 0 deg. and 26.4 deg. Trailing-edge flaps were deflected 0 deg., 10 deg., 12.9 deg., and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein. The data associated with each deflected leading-edge flap indicate L/D improvements over the undeflected configuration. These improvements may be instrumental in providing the necessary lift augmentation required by an actual HSCT during the climb-out and landing phases of the flight envelope. However, further tests will have to be done to assess their full potential.

Efficiency enhancement of a self-propelled pitching profile using non-sinusoidal trajectories

NASA Astrophysics Data System (ADS)

Mekadem, M.; Chihani, E.; Oualli, H.; Hanchi, S.; Bouabdallah, A.; Gad-El-Hak, M.

2017-11-01

A symmetrical profile is subjected to non-sinusoidal pitching motion. The airfoil has a chord length c = 0.006 m and a semi-circular leading edge with a diameter of D = 0.001 m. The extrados and intrados are two straight lines that intersect at a tapered trailing edge, and the pitching pivot point is positioned at the leading edge. The pitching frequency is in the range of 1 <= f <= 190 Hz, while the tangential amplitude of the flapping trailing edge varies from 18% to 114% of the foil cord. To improve the airfoil propulsive performance, two-dimensional numerical simulations are implemented on FLUENT. The Reynolds number based upon the maximum profile thickness D varies in the range of 35 <= Re <= 210 , which matches insect's Reynolds numbers. The foil movement is executed using the dynamic mesh technique and a user defined function (UDF). The adopted mesh has 70,445 nodes with 5,1960 quadrilateral cells. The results are in good agreement with prior experiments, and, compared to sinusoidal oscillations, show that non-sinusoidal flapping trajectories lead to advancing velocity increase of 550%. Additionally, if improved propulsive efficiency is sought, non-sinusoidal flapping lead to better thrust.

Experimental Test Results of Energy Efficient Transport (EET) High-Lift Airfoil in Langley Low-Turbulence Pressure Tunnel

NASA Technical Reports Server (NTRS)

Morgan, Harry L., Jr.

2002-01-01

This report describes the results of an experimental study conducted in the Langley Low-Turbulence Pressure Tunnel to determine the effects of Reynolds number and Mach number on the two-dimensional aerodynamic performance of the Langley Energy Efficient Transport (EET) High-Lift Airfoil. The high-lift airfoil was a supercritical-type airfoil with a thickness-to- chord ratio of 0.12 and was equipped with a leading-edge slat and a double-slotted trailing-edge flap. The leading-edge slat could be deflected -30 deg, -40 deg, -50 deg, and -60 deg, and the trailing-edge flaps could be deflected to 15 deg, 30 deg, 45 deg, and 60 deg. The gaps and overlaps for the slat and flaps were fixed at each deflection resulting in 16 different configurations. All 16 configurations were tested through a Reynolds number range of 2.5 to 18 million at a Mach number of 0.20. Selected configurations were also tested through a Mach number range of 0.10 to 0.35. The plotted and tabulated force, moment, and pressure data are available on the CD-ROM supplement L-18221.

Low-speed aerodynamic characteristics from wind-tunnel tests of a large-scale advanced arrow-wing supersonic-cruise transport concept

NASA Technical Reports Server (NTRS)

Smith, P. M.

1978-01-01

Tests have been conducted to extend the existing low speed aerodynamic data base of advanced supersonic-cruise arrow wing configurations. Principle configuration variables included wing leading-edge flap deflection, wing trailing-edge flap deflection, horizontal tail effectiveness, and fuselage forebody strakes. A limited investigation was also conducted to determine the low speed aerodynamic effects due to slotted training-edge flaps. Results of this investigation demonstrate that deflecting the wing leading-edge flaps downward to suppress the wing apex vortices provides improved static longitudinal stability; however, it also results in significantly reduced static directional stability. The use of a selected fuselage forebody strakes is found to be effective in increasing the level of positive static directional stability. Drooping the fuselage nose, which is required for low-speed pilot vision, significantly improves the later-directional trim characteristics.

Lift augmentation on a delta wing via leading edge fences and the Gurney flap. M.S. Thesis

NASA Technical Reports Server (NTRS)

Buchholz, Mark D.

1992-01-01

Wind tunnel tests were conducted on two devices for the purpose of lift augmentation on a 60 deg delta wing at low speed. Lift, drag, pitching moment, and surface pressures were measured. Detailed flow visualization was also obtained. Both the leading edge fence and the Gurney flap are shown to increase lift. The fences and flap shift the lift curve as much as 5 deg and 10 deg, respectively. The fences aid in trapping vortices on the upper surface, thereby increasing suction. The Gurney flap improves circulation at the trailing edge. The individual influences of both devices are roughly additive, creating high lift gain. However, the lower lift to drag ratio and the precipitation of vortex burst caused by the fences, and the nose down pitching moment created by the flap are also significant factors.

USB noise reduction by nozzle and flap modifications

NASA Technical Reports Server (NTRS)

Hayden, R. E.

1976-01-01

The development of concepts for reducing upper surface blown flap noise at the source through flap modifications and special nozzles is reviewed. In particular, recent results obtained on the aerodynamic and acoustic performance of flaps with porous surfaces near the trailing edge and multi-slotted nozzles are reviewed. Considerable reduction (6-10 db) of the characteristic low frequency peak is shown. The aerodynamic performance is compared with conventional systems, and prospects for future improvements are discussed.

The effects of gusts on the fluctuating airloads of airfoils in transonic flow

NASA Technical Reports Server (NTRS)

Mccroskey, W. J.

1984-01-01

Unsteady interactions of distributed and sharp-edged gusts with a stationary airfoil have been analyzed in two-dimensional transonic flow.A simple method of introducing such disturbances has been numerically implemented within the framework of unsteady, transonic small-disturbance theory. Representative solutions for various airfoils subjected to chordwise and transverse gusts show that the strength and unsteady motion of the shock wave on the airfoil significantly affect the flowfield development and, consequently, the dynamic airloads. Also a study was made of the reductions in the unsteady airloads that can be achieved by the proper active control motion of a trailing-edge flap, and a simple gust-alleviation strategy was developed. However, the chordwise pressure distributions associated with gusts are very different from those produced by trailing-edge flap oscillations. Consequently, the fluctuating lift and the unsteady pitching moments cannot both be eliminated simultaneously.

Applicability of linearized-theory attached-flow methods to design and analysis of flap systems at low speeds for thin swept wings with sharp leading edges

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.

1987-01-01

Low-speed experimental force and data on a series of thin swept wings with sharp leading edges and leading and trailing-edge flaps are compared with predictions made using a linearized-theory method which includes estimates of vortex forces. These comparisons were made to assess the effectiveness of linearized-theory methods for use in the design and analysis of flap systems in subsonic flow. Results demonstrate that linearized-theory, attached-flow methods (with approximate representation of vortex forces) can form the basis of a rational system for flap design and analysis. Even attached-flow methods that do not take vortex forces into account can be used for the selection of optimized flap-system geometry, but design-point performance levels tend to be underestimated unless vortex forces are included. Illustrative examples of the use of these methods in the design of efficient low-speed flap systems are included.

Advancements in adaptive aerodynamic technologies for airfoils and wings

NASA Astrophysics Data System (ADS)

Jepson, Jeffrey Keith

Although aircraft operate over a wide range of flight conditions, current fixed-geometry aircraft are optimized for only a few of these conditions. By altering the shape of the aircraft, adaptive aerodynamics can be used to increase the safety and performance of an aircraft by tailoring the aircraft for multiple flight conditions. Of the various shape adaptation concepts currently being studied, the use of multiple trailing-edge flaps along the span of a wing offers a relatively high possibility of being incorporated on aircraft in the near future. Multiple trailing-edge flaps allow for effective spanwise camber adaptation with resulting drag benefits over a large speed range and load alleviation at high-g conditions. The research presented in this dissertation focuses on the development of this concept of using trailing-edge flaps to tailor an aircraft for multiple flight conditions. One of the major tasks involved in implementing trailing-edge flaps is in designing the airfoil to incorporate the flap. The first part of this dissertation presents a design formulation that incorporates aircraft performance considerations in the inverse design of low-speed laminar-flow adaptive airfoils with trailing-edge cruise flaps. The benefit of using adaptive airfoils is that the size of the low-drag region of the drag polar can be effectively increased without increasing the maximum thickness of the airfoil. Two aircraft performance parameters are considered: level-flight maximum speed and maximum range. It is shown that the lift coefficients for the lower and upper corners of the airfoil low-drag range can be appropriately adjusted to tailor the airfoil for these two aircraft performance parameters. The design problem is posed as a part of a multidimensional Newton iteration in an existing conformal-mapping based inverse design code, PROFOIL. This formulation automatically adjusts the lift coefficients for the corners of the low-drag range for a given flap deflection as required for the airfoil-aircraft matching. Examples are presented to illustrate the flapped-airfoil design approach for a general aviation aircraft and the results are validated by comparison with results from post-design aircraft performance computations. Once the airfoil is designed to incorporate a TE flap, it is important to determine the most suitable flap angles along the wing for different flight conditions. The second part of this dissertation presents a method for determining the optimum flap angles to minimize drag based on pressures measured at select locations on the wing. Computational flow simulations using a panel method are used "in the loop" for demonstrating closed-loop control of the flaps. Examples in the paper show that the control algorithm is successful in correctly adapting the wing to achieve the target lift distributions for minimizing induced drag while adjusting the wing angle of attack for operation of the wing in the drag bucket. It is shown that the "sense-and-adapt" approach developed is capable of handling varying and unpredictable inflow conditions. Such a capability could be useful in adapting long-span flexible wings that may experience significant and unknown atmospheric inflow variations along the span. To further develop the "sense-and-adapt" approach, the method was tested experimentally in the third part of the research. The goal of the testing was to see if the same results found computationally can be obtained experimentally. The North Carolina State University subsonic wind tunnel was used for the wind tunnel tests. Results from the testing showed that the "sense-and-adapt" approach has the same performance experimentally as it did computationally. The research presented in this dissertation is a stepping stone towards further development of the concept, which includes modeling the system in the Simulink environment and flight experiments using uninhabited aerial vehicles.

Flap survey test of a combined surface blowing model: Flow measurements at static flow conditions

NASA Technical Reports Server (NTRS)

Fukushima, T.

1978-01-01

The Combined Surface Blowing (CSB) V/STOL lift/propulsion system consists of a blown flap system which deflects the exhaust from a turbojet engine over a system of flaps deployed at the trailing edge of the wing. Flow measurements consisting of velocity measurements using split film probes and total measure surveys using a miniature Kiel probe were made at control stations along the flap systems at two spanwise stations, the centerline of the nozzle and 60 percent of the nozzle span outboard of the centerline. Surface pressure measurements were made in the wing cove and the upper surface of the first flap element. The test showed a significant flow separation in the wing cove. The extent of the separation is so large that the flow into the first flap takes place only at the leading edge of the flap. The velocity profile measurements indicate that large spanwise (3 dimensional) flow may exist.

Low-Speed Wind-Tunnel Investigation to Determine the Aerodynamic Characteristics of a Rectangular Wing Equipped with a Full-Span and an Inboard Half-Span Jet-Augmented Flap Deflected 55 deg

NASA Technical Reports Server (NTRS)

Gainer, Thomas G.

1959-01-01

An investigation to determine the aerodynamic characteristics of a rectangular wing equipped with a full-span and an inboard half-span jet-augmented flap has been made in the Langley 300 MPH 7- by 10-foot tunnel. The wing had an aspect ratio of 8.3 and a thickness-chord ratio of 0.167. A jet of air was blown backward through a small gap, tangentially to the upper surface of a round trailing edge, and was separated from the trailing edge by a very small flap at an angle of 55 deg with respect to the wing-chord plane. The results of the investigation showed that the ratio of total lift to jet-reaction lift for the wing was about 35 percent less for the half-span jet-augmented flap than for the full-span jet-augmented flap. The reduction of the span of the jet-augmented flap from full to half span reduced the maximum value of jet-circulation lift coefficient that could be produced from about 6.8 to a value of about 2.2. The half-span jet- augmented flap gave thrust recoveries considerably poorer than those obtained with the full-span jet-augmented flap. Large nose-down pitching- moment coefficients were produced by the half-span flap, with the greater part of these being the result of the larger jet reactions required to produce a given lift for the half-spin flap compared with that required for the full-span flap.

AERO2S - SUBSONIC AERODYNAMIC ANALYSIS OF WINGS WITH LEADING- AND TRAILING-EDGE FLAPS IN COMBINATION WITH CANARD OR HORIZONTAL TAIL SURFACES (IBM PC VERSION)

NASA Technical Reports Server (NTRS)

Carlson, H. W.

1994-01-01

This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is necessary only to add an identification record and the namelist data that are to be changed from the previous run. This code was originally developed in 1989 in FORTRAN V on a CDC 6000 computer system, and was later ported to an MS-DOS environment. Both versions are available from COSMIC. There are only a few differences between the PC version (LAR-14458) and CDC version (LAR-14178) of AERO2S distributed by COSMIC. The CDC version has one main source code file while the PC version has two files which are easier to edit and compile on a PC. The PC version does not require a FORTRAN compiler which supports NAMELIST because a special INPUT subroutine has been added. The CDC version includes two MODIFY decks which can be used to improve the code and prevent the possibility of some infrequently occurring errors while PC-version users will have to make these code changes manually. The PC version includes an executable which was generated with the Ryan McFarland/FORTRAN compiler and requires 253K RAM and an 80x87 math co-processor. Using this executable, the sample case requires about four hours to execute on an 8MHz AT-class microcomputer with a co-processor. The source code conforms to the FORTRAN 77 standard except that it uses variables longer than six characters. With two minor modifications, the PC version should be portable to any computer with a FORTRAN compiler and sufficient memory. The CDC version of AERO2S is available in CDC NOS Internal format on a 9-track 1600 BPI magnetic tape. The PC version is available on a set of two 5.25 inch 360K MS-DOS format diskettes. IBM AT is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. CDC is a registered trademark of Control Data Corporation. NOS is a trademark of Control Data Corporation.

AERO2S - SUBSONIC AERODYNAMIC ANALYSIS OF WINGS WITH LEADING- AND TRAILING-EDGE FLAPS IN COMBINATION WITH CANARD OR HORIZONTAL TAIL SURFACES (CDC VERSION)

NASA Technical Reports Server (NTRS)

Darden, C. M.

1994-01-01

This code was developed to aid design engineers in the selection and evaluation of aerodynamically efficient wing-canard and wing-horizontal-tail configurations that may employ simple hinged-flap systems. Rapid estimates of the longitudinal aerodynamic characteristics of conceptual airplane lifting surface arrangements are provided. The method is particularly well suited to configurations which, because of high speed flight requirements, must employ thin wings with highly swept leading edges. The code is applicable to wings with either sharp or rounded leading edges. The code provides theoretical pressure distributions over the wing, the canard or horizontal tail, and the deflected flap surfaces as well as estimates of the wing lift, drag, and pitching moments which account for attainable leading edge thrust and leading edge separation vortex forces. The wing planform information is specified by a series of leading edge and trailing edge breakpoints for a right hand wing panel. Up to 21 pairs of coordinates may be used to describe both the leading edge and the trailing edge. The code has been written to accommodate 2000 right hand panel elements, but can easily be modified to accommodate a larger or smaller number of elements depending on the capacity of the target computer platform. The code provides solutions for wing surfaces composed of all possible combinations of leading edge and trailing edge flap settings provided by the original deflection multipliers and by the flap deflection multipliers. Up to 25 pairs of leading edge and trailing edge flap deflection schedules may thus be treated simultaneously. The code also provides for an improved accounting of hinge-line singularities in determination of wing forces and moments. To determine lifting surface perturbation velocity distributions, the code provides for a maximum of 70 iterations. The program is constructed so that successive runs may be made with a given code entry. To make additional runs, it is necessary only to add an identification record and the namelist data that are to be changed from the previous run. This code was originally developed in 1989 in FORTRAN V on a CDC 6000 computer system, and was later ported to an MS-DOS environment. Both versions are available from COSMIC. There are only a few differences between the PC version (LAR-14458) and CDC version (LAR-14178) of AERO2S distributed by COSMIC. The CDC version has one main source code file while the PC version has two files which are easier to edit and compile on a PC. The PC version does not require a FORTRAN compiler which supports NAMELIST because a special INPUT subroutine has been added. The CDC version includes two MODIFY decks which can be used to improve the code and prevent the possibility of some infrequently occurring errors while PC-version users will have to make these code changes manually. The PC version includes an executable which was generated with the Ryan McFarland/FORTRAN compiler and requires 253K RAM and an 80x87 math co-processor. Using this executable, the sample case requires about four hours to execute on an 8MHz AT-class microcomputer with a co-processor. The source code conforms to the FORTRAN 77 standard except that it uses variables longer than six characters. With two minor modifications, the PC version should be portable to any computer with a FORTRAN compiler and sufficient memory. The CDC version of AERO2S is available in CDC NOS Internal format on a 9-track 1600 BPI magnetic tape. The PC version is available on a set of two 5.25 inch 360K MS-DOS format diskettes. IBM AT is a registered trademark of International Business Machines. MS-DOS is a registered trademark of Microsoft Corporation. CDC is a registered trademark of Control Data Corporation. NOS is a trademark of Control Data Corporation.

Factors Influencing the Accuracy of Aerodynamic Hinge-Moment Prediction

DTIC Science & Technology

1978-08-01

condition on the aft lifting surfaces and flaps. A new modeling technique for trailing-edge wake analysis using a potential- flow program based on the...control surface as depicLed in figure 21.. Three different models are used to simulate the flow on the wing, the flap, and the gaps. In the first two panel...ized sense, similar to that implemented in the FLEXSTAB program. The modeling of the wake on the side-edge gaps differs in the first two panel models

Chordwise implementation of pneumatic artificial muscles to actuate a trailing edge flap

NASA Astrophysics Data System (ADS)

Vocke, R. D., III; Kothera, C. S.; Wereley, N. M.

2018-07-01

This work describes the theoretical design and experimental validation of a rotorcraft-specific trailing edge flap powered by pneumatic artificial muscle actuators. The actuators in this work are co-located outboard on the rotor blade with the flap and arranged with a chordwise orientation where diameter and length restrictions can severely limit the operating range of the system. Techniques for addressing this configuration, such as introducing a bias contraction and mechanism optimization, are discussed and a numerical optimization is performed for an actuation system sized for implementation on a medium utility helicopter rotor. The optimized design achieves ±10° of deflection at 1/rev, and maintains at least ±2° half peak-to-peak deflection out to 10/rev, indicating that the system has the actuation authority and bandwidth necessary for both primary control and vibration/noise reduction. Portions of this paper were presented at the AHS 70th Annual Forum, Montréal, Québec, Canada, May 20–22, 2014.

Wind-tunnel tests on model wing with Fowler flap and specially developed leading-edge slot

NASA Technical Reports Server (NTRS)

Weick, Fred E; Platt, Robert C

1933-01-01

An investigation was made in the NACA 7 by 10 foot wind tunnel to find the increase in maximum lift coefficient which could be obtained by providing a model wing with both a Fowler trailing-edge extension flap and a Handley Page type leading-edge slot. A conventional Handley page slot proportioned to operate on the plain wing without a flap gave but a slight increase with the flap; so a special form of slot was developed to work more effectively with the flap. With the best combined arrangement the maximum lift coefficient based on the original area was increased from 3.17, for the Fowler wing, to 3.62. The minimum drag coefficient with both devices retracted was increased in approximately the same proportion. Tests were also made with the special-type slot on the plain wing without the flap. The special slot, used either with or without the Fowler flap, gave definitely higher values of the maximum lift coefficient than the slots of conventional form, with an increase of the same order in the minimum drag coefficient.

KSC-04pd0464

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - The body flap is installed on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0463

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - The body flap is installed on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

USB flow characteristics related to noise generation

NASA Technical Reports Server (NTRS)

Brown, W. H.; Reddy, N. N.

1976-01-01

The effects of nozzle and flap geometry on upper surface blown flow field characteristics related to noise generation were examined experimentally using static models. Flow attachment and spreading characteristics were observed using flow visualization techniques. Velocity and turbulence profiles in the trailing edge wake were measured using hot-wire anemometry, and the effects of the geometric variables on peak velocity and turbulence intensity were determined. It is shown that peak trailing edge velocity is a function of the ratio of flow length to modified hydraulic diameter.

Spin-Tunnel Investigation of a 1/20-Scale Model of the Northrop F-5E Airplane

NASA Technical Reports Server (NTRS)

Scher, Stanley H.; White, William L.

1977-01-01

An investigation has been conducted in the Langley spin tunnel to determine the spin and recovery characteristics of a 1/20-scale model of the Northrop F-5E airplane. The investigation included erect and inverted spins, a range of center-of- gravity locations and moments of inertia, symmetric and asymmetric store loadings, and a determination of the parachute size required for emergency spin recovery. The effects of increased elevator trailing-edge-up deflections, of leading-edge and trailing-edge flap deflections, and of simulating the geometry of large external stores were also determined.

Active Tailoring of Lift Distribution to Enhance Cruise Performance

NASA Technical Reports Server (NTRS)

Flamm, Jeffrey D. (Technical Monitor); Pfeiffer, Neal J.; Christians, Joel G.

2005-01-01

During Phase I of this project, Raytheon Aircraft Company (RAC) has analytically and experimentally evaluated key components of a system that could be implemented for active tailoring of wing lift distribution using low-drag, trailing-edge modifications. Simple systems such as those studied by RAC could be used to enhance the cruise performance of a business jet configuration over a range of typical flight conditions. The trailing-edge modifications focus on simple, deployable mechanisms comprised of extendable small flap panels over portions of the span that could be used to subtly but positively optimize the lift and drag characteristics. The report includes results from low speed wind tunnel testing of the trailing-edge devices, descriptions of potential mechanisms for automation, and an assessment of the technology.

Characteristics of merging shear layers and turbulent wakes of a multi-element airfoil

NASA Technical Reports Server (NTRS)

Adair, Desmond; Horne, W. Clifton

1988-01-01

Flow characteristics in the vicinity of the trailing edge of a single-slotted airfoil flap are presented and analyzed. The experimental arrangement consisted of a NACA 4412 airfoil equipped with a NACA 4415 flap whose angle of deflection was 21.8 deg. The flow remained attached over the model surfaces except in the vicinity of the flap trailing edge where a small region of boundary-layer separation extended over the aft 7 percent of flap chord. The flow was complicated by the presence of a strong, initially inviscid jet emanating from the slot between airfoil and flap, and a gradual merging of the main airfoil wake and flap suction-side boundary layer. Downstream of the flap, the airfoil and flap wakes fully merged to form an asymmetrical curved wake. The airfoil configuration was tested at an angle of attack of 8.2 deg, at a Mach number of 0.09, and a chord based Reynolds number of 1.8 x 10 to the 6th power in the Ames Research Center 7- by 10-Foot Wind Tunnel. Surface pressure measurements were made on the airfoil and flap and on the wind tunnel roof and floor. It was estimated that the wall interference increased the C sub L by 7 percent and decreased the C sub M by 4.5 percent. Velocity characteristics were quantified using hot-wire anemometry in regions of flow with preferred direction and low turbulence intensity. A 3-D laser velocimeter was used in regions of flow recirculation and relatively high turbulence intensity.

DETAIL OF UNDERSIDE OF RIGHT WING. AFTER DOCKING, THE FIRST ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

DETAIL OF UNDERSIDE OF RIGHT WING. AFTER DOCKING, THE FIRST STEP IN MAINTENANCE IS TO OPEN UP THOSE AREAS OF THE AIRCRAFT THAT REQUIRE SERVICING. IN THIS VIEW WING SLATS, KRUGER FLAPS, AND TRAILING EDGE FLAPS ARE ALL DOWN; THRUST REVERSERS ARE IN THE OPEN POSITION; ENGINE COWLING IS OPEN. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

Vibration reduction in helicopter rotors using an actively controlled partial span trailing edge flap located on the blade

NASA Technical Reports Server (NTRS)

Millott, T. A.; Friedmann, P. P.

1994-01-01

This report describes an analytical study of vibration reduction in a four-bladed helicopter rotor using an actively controlled, partial span, trailing edge flap located on the blade. The vibration reduction produced by the actively controlled flap (ACF) is compared with that obtained using individual blade control (IBC), in which the entire blade is oscillated in pitch. For both cases a deterministic feedback controller is implemented to reduce the 4/rev hub loads. For all cases considered, the ACF produced vibration reduction comparable with that obtained using IBC, but consumed only 10-30% of the power required to implement IBC. A careful parametric study is conducted to determine the influence of blade torsional stiffness, spanwise location of the control flap, and hinge moment correction on the vibration reduction characteristics of the ACF. The results clearly demonstrate the feasibility of this new approach to vibration reduction. It should be emphasized than the ACF, used together with a conventional swashplate, is completely decoupled from the primary flight control system and thus it has no influence on the airworthiness of the helicopter. This attribute is potentially a significant advantage when compared to IBC.

Small scale model static acoustic investigation of hybrid high lift systems combining upper surface blowing with the internally blown flap

NASA Technical Reports Server (NTRS)

Cole, T. W.; Rathburn, E. A.

1974-01-01

A static acoustic and propulsion test of a small radius Jacobs-Hurkamp and a large radius Flex Flap combined with four upper surface blowing (USB) nozzles was performed. Nozzle force and flow data, flap trailing edge total pressure survey data, and acoustic data were obtained. Jacobs-Hurkamp flap surface pressure data, flow visualization photographs, and spoiler acoustic data from the limited mid-year tests are reported. A pressure ratio range of 1.2 to 1.5 was investigated for the USB nozzles and for the auxiliary blowing slots. The acoustic data were scaled to a four-engine STOL airplane of roughly 110,000 kilograms or 50,000 pounds gross weight, corresponding to a model scale of approximately 0.2 for the nozzles without deflector. The model nozzle scale is actually reduced to about .17 with deflector although all results in this report assume 0.2 scale factor. Trailing edge pressure surveys indicated that poor flow attachment was obtained even at large flow impingement angles unless a nozzle deflector plate was used. Good attachment was obtained with the aspect ratio four nozzle with deflector, confirming the small scale wind tunnel tests.

Full-scale-wind-tunnel Tests of a 35 Degree Sweptback Wing Airplane with High-velocity Blowing over the Training-edge Flaps

NASA Technical Reports Server (NTRS)

Kelley, Mark W; Tolhurst, William H JR

1955-01-01

A wind-tunnel investigation was made to determine the effects of ejecting high-velocity air near the leading edge of plain trailing-edge flaps on a 35 degree sweptback wing. The tests were made with flap deflections from 45 degrees to 85 degrees and with pressure ratios across the flap nozzles from sub-critical up to 2.9. A limited study of the effects of nozzle location and configuration on the efficiency of the flap was made. Measurements of the lift, drag, and pitching moment were made for Reynolds numbers from 5.8 to 10.1x10(6). Measurements were also made of the weight rate of flow, pressure, and temperature of the air supplied to the flap nozzles.The results show that blowing on the deflected flap produced large flap lift increments. The amount of air required to prevent flow separation on the flap was significantly less than that estimated from published two-dimensional data. When the amount of air ejected over the flap was just sufficient to prevent flow separation, the lift increment obtained agreed well with linear inviscid fluid theory up to flap deflections of 60 degrees. The flap lift increment at 85 degrees flap deflection was about 80 percent of that predicted theoretically.With larger amounts of air blown over the flap, these lift increments could be significantly increased. It was found that the performance of the flap was relatively insensitive to the location of the flap nozzle, to spacers in the nozzle, and to flow disturbances such as those caused by leading-edge slats or discontinuities on the wing or flap surfaces. Analysis of the results indicated that installation of this system on an F-86 airplane is feasible.

Wind tunnel investigation of a high lift system with pneumatic flow control

NASA Astrophysics Data System (ADS)

Victor, Pricop Mihai; Mircea, Boscoianu; Daniel-Eugeniu, Crunteanu

2016-06-01

Next generation passenger aircrafts require more efficient high lift systems under size and mass constraints, to achieve more fuel efficiency. This can be obtained in various ways: to improve/maintain aerodynamic performance while simplifying the mechanical design of the high lift system going to a single slotted flap, to maintain complexity and improve the aerodynamics even more, etc. Laminar wings have less efficient leading edge high lift systems if any, requiring more performance from the trailing edge flap. Pulsed blowing active flow control (AFC) in the gap of single element flap is investigated for a relatively large model. A wind tunnel model, test campaign and results and conclusion are presented.

Variable Camber Continuous Aerodynamic Control Surfaces and Methods for Active Wing Shaping Control

NASA Technical Reports Server (NTRS)

Nguyen, Nhan T. (Inventor)

2016-01-01

An aerodynamic control apparatus for an air vehicle improves various aerodynamic performance metrics by employing multiple spanwise flap segments that jointly form a continuous or a piecewise continuous trailing edge to minimize drag induced by lift or vortices. At least one of the multiple spanwise flap segments includes a variable camber flap subsystem having multiple chordwise flap segments that may be independently actuated. Some embodiments also employ a continuous leading edge slat system that includes multiple spanwise slat segments, each of which has one or more chordwise slat segment. A method and an apparatus for implementing active control of a wing shape are also described and include the determination of desired lift distribution to determine the improved aerodynamic deflection of the wings. Flap deflections are determined and control signals are generated to actively control the wing shape to approximate the desired deflection.

Measurements of noise produced by flow past lifting surfaces

NASA Technical Reports Server (NTRS)

Kendall, J. M.

1978-01-01

Wind tunnel studies have been conducted to determine the specific locations of aerodynamic noise production within the flow field about various lifting-surface configurations. The models tested included low aspect ratio shapes intended to represent aircraft flaps, a finite aspect ratio NACA 0012 wing, and a multi-element wing section consisting of a main section, a leading edge flap, and dual trailing edge flaps. Turbulence was induced on the models by surface roughness. Lift and drag were measured for the flap models. Hot-wire anemometry was used for study of the flap-model vortex roll-up. Apparent noise source distributions were measured by use of a directional microphone system, located outside the tunnel, which was scanned about the flow region to be analyzed under computer control. These distributions exhibited a diversity of pattern, suggesting that several flow processes are important to lifting-surface noise production. Speculation concerning these processes is offered.

KSC-04pd0459

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility help move the body flap into position on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0461

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - A worker on a ladder (lower left) observes installation of the body flap onto the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0462

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - Workers on ladders (left and right) check installation of the body flap onto the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0452

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the body flap for the orbiter Discovery is prepared for installation. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04PD-0462

NASA Technical Reports Server (NTRS)

2004-01-01

KENNEDY SPACE CENTER, FLA. Workers on ladders (left and right) check installation of the body flap onto the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

Vector thrust induced lift effects for several ejector exhaust locations on a V/STOL wind tunnel model at forward speed

NASA Technical Reports Server (NTRS)

Sharon, A. D.

1975-01-01

The results and analysis of aerodynamic force data obtained from a small scale model of a V/STOL research vehicle in a low speed wind tunnel are presented. The analysis of the data includes the evaluation of aerodynamic-propulsive lift performance when operating twin ejector nozzles with thrust deflected. Three different types of thrust deflector systems were examined: 90 deg downward deflected nozzle, 90 deg slotted nozzle with boundary layer control, and an externally blown flap configuration. Several nozzle locations were tested, including over and underwing positions. The interference lift of the nacelle and model due to jet exhaust thrust is compared and results show that 90 deg turned nozzles located over the wing (near the trailing edge) produce the largest interference lift increment for an untrimmed aircraft, and that the slotted nozzle located under the wing near the trailing edge (in conjunction with a BLC flap) gives a comparable interference lift in the trimmed condition. The externally blown flap nozzle produced the least interference lift and significantly less total lift due to jet thrust effects.

Characterization of Flap Edge Noise Radiation from a High-Fidelity Airframe Model

NASA Technical Reports Server (NTRS)

Humphreys, William M., Jr.; Khorrami, Mehdi R.; Lockard, David P.; Neuhart, Dan H.; Bahr, Christopher J.

2015-01-01

The results of an experimental study of the noise generated by a baseline high-fidelity airframe model are presented. The test campaign was conducted in the open-jet test section of the NASA Langley 14- by 22-foot Subsonic Tunnel on an 18%-scale, semi-span Gulfstream airframe model incorporating a trailing edge flap and main landing gear. Unsteady surface pressure measurements were obtained from a series of sensors positioned along the two flap edges, and far field acoustic measurements were obtained using a 97-microphone phased array that viewed the pressure side of the airframe. The DAMAS array deconvolution method was employed to determine the locations and strengths of relevant noise sources in the vicinity of the flap edges and the landing gear. A Coherent Output Power (COP) spectral method was used to couple the unsteady surface pressures measured along the flap edges with the phased array output. The results indicate that outboard flap edge noise is dominated by the flap bulb seal cavity with very strong COP coherence over an approximate model-scale frequency range of 1 to 5 kHz observed between the array output and those unsteady pressure sensors nearest the aft end of the cavity. An examination of experimental COP spectra for the inboard flap proved inconclusive, most likely due to a combination of coherence loss caused by decorrelation of acoustic waves propagating through the thick wind tunnel shear layer and contamination of the spectra by tunnel background noise at lower frequencies. Directivity measurements obtained from integration of DAMAS pressure-squared values over defined geometric zones around the model show that the baseline flap and landing gear are only moderately directional as a function of polar emission angle.

Computational wing optimization and comparisons with experiment for a semi-span wing model

NASA Technical Reports Server (NTRS)

Waggoner, E. G.; Haney, H. P.; Ballhaus, W. F.

1978-01-01

A computational wing optimization procedure was developed and verified by an experimental investigation of a semi-span variable camber wing model in the NASA Ames Research Center 14 foot transonic wind tunnel. The Bailey-Ballhaus transonic potential flow analysis and Woodward-Carmichael linear theory codes were linked to Vanderplaats constrained minimization routine to optimize model configurations at several subsonic and transonic design points. The 35 deg swept wing is characterized by multi-segmented leading and trailing edge flaps whose hinge lines are swept relative to the leading and trailing edges of the wing. By varying deflection angles of the flap segments, camber and twist distribution can be optimized for different design conditions. Results indicate that numerical optimization can be both an effective and efficient design tool. The optimized configurations had as good or better lift to drag ratios at the design points as the best designs previously tested during an extensive parametric study.

Continuous Trailing-Edge Flaps for Primary Flight Control of a Helicopter Main Rotor

NASA Technical Reports Server (NTRS)

Thornburgh, Robert P.; Kreshock, Andrew R.; Wilbur, Matthew L.; Sekula, Martin K.; Shen, Jinwei

2014-01-01

The use of continuous trailing-edge flaps (CTEFs) for primary flight control of a helicopter main rotor is studied. A practical, optimized bimorph design with Macro-Fiber Composite actuators is developed for CTEF control, and a coupled structures and computational fluid dynamics methodology is used to study the fundamental behavior of an airfoil with CTEFs. These results are used within a comprehensive rotorcraft analysis model to study the control authority requirements of the CTEFs when utilized for primary flight control of a utility class helicopter. A study of the effect of blade root pitch index (RPI) on CTEF control authority is conducted, and the impact of structural and aerodynamic model complexity on the comprehensive analysis results is presented. The results show that primary flight control using CTEFs is promising; however, a more viable option may include the control of blade RPI, as well.

Lift enhancing tabs for airfoils

NASA Technical Reports Server (NTRS)

Ross, James C. (Inventor)

1994-01-01

A tab deployable from the trailing edge of a main airfoil element forces flow onto a following airfoil element, such as a flap, to keep the flow attached and thus enhance lift. For aircraft wings with high lift systems that include leading edge slats, the slats may also be provided with tabs to turn the flow onto the following main element.

Allometry of wing twist and camber in a flower chafer during free flight: How do wing deformations scale with body size?

PubMed Central

Ribak, Gal

2017-01-01

Intraspecific variation in adult body mass can be particularly high in some insect species, mandating adjustment of the wing's structural properties to support the weight of the larger body mass in air. Insect wings elastically deform during flapping, dynamically changing the twist and camber of the relatively thin and flat aerofoil. We examined how wing deformations during free flight scale with body mass within a species of rose chafers (Coleoptera: Protaetia cuprea) in which individuals varied more than threefold in body mass (0.38–1.29 g). Beetles taking off voluntarily were filmed using three high-speed cameras and the instantaneous deformation of their wings during the flapping cycle was analysed. Flapping frequency decreased in larger beetles but, otherwise, flapping kinematics remained similar in both small and large beetles. Deflection of the wing chord-wise varied along the span, with average deflections at the proximal trailing edge higher by 0.2 and 0.197 wing lengths compared to the distal trailing edge in the downstroke and the upstroke, respectively. These deflections scaled with wing chord to the power of 1.0, implying a constant twist and camber despite the variations in wing and body size. This suggests that the allometric growth in wing size includes adjustment of the flexural stiffness of the wing structure to preserve wing twist and camber during flapping. PMID:29134103

Allometry of wing twist and camber in a flower chafer during free flight: How do wing deformations scale with body size?

PubMed

Meresman, Yonatan; Ribak, Gal

2017-10-01

Intraspecific variation in adult body mass can be particularly high in some insect species, mandating adjustment of the wing's structural properties to support the weight of the larger body mass in air. Insect wings elastically deform during flapping, dynamically changing the twist and camber of the relatively thin and flat aerofoil. We examined how wing deformations during free flight scale with body mass within a species of rose chafers (Coleoptera: Protaetia cuprea ) in which individuals varied more than threefold in body mass (0.38-1.29 g). Beetles taking off voluntarily were filmed using three high-speed cameras and the instantaneous deformation of their wings during the flapping cycle was analysed. Flapping frequency decreased in larger beetles but, otherwise, flapping kinematics remained similar in both small and large beetles. Deflection of the wing chord-wise varied along the span, with average deflections at the proximal trailing edge higher by 0.2 and 0.197 wing lengths compared to the distal trailing edge in the downstroke and the upstroke, respectively. These deflections scaled with wing chord to the power of 1.0, implying a constant twist and camber despite the variations in wing and body size. This suggests that the allometric growth in wing size includes adjustment of the flexural stiffness of the wing structure to preserve wing twist and camber during flapping.

Noise Reduction Through Circulation Control

NASA Technical Reports Server (NTRS)

Munro, Scott E.; Ahuja, K. K.; Englar, Robert J.

2005-01-01

Circulation control technology uses tangential blowing around a rounded trailing edge or a leading edge to change the force and moment characteristics of an aerodynamic body. This technology has been applied to circular cylinders, wings, helicopter rotors, and even to automobiles for improved aerodynamic performance. Only limited research has been conducted on the acoustic of this technology. Since wing flaps contribute to the environmental noise of an aircraft, an alternate blown high lift system without complex mechanical flaps could prove beneficial in reducing the noise of an approaching aircraft. Thus, in this study, a direct comparison of the acoustic characteristics of high lift systems employing a circulation control wing configuration and a conventional wing flapped configuration has been made. These results indicate that acoustically, a circulation control wing high lift system could be considerably more acceptable than a wing with conventional mechanical flaps.

ACTE Wing Loads Analysis

NASA Technical Reports Server (NTRS)

Horn, Nicholas R.

2015-01-01

The Adaptive Compliant Trailing Edge (ACTE) project modified a Gulfstream III (GIII) aircraft with a new flexible flap that creates a seamless transition between the flap and the wing. As with any new modification, it is crucial to ensure that the aircraft will not become overstressed in flight. To test this, Star CCM a computational fluid dynamics (CFD) software program was used to calculate aerodynamic data for the aircraft at given flight conditions.

AFTI/F-111 MAW flight control system and redundancy management description

NASA Technical Reports Server (NTRS)

Larson, Richard R.

1987-01-01

The wing on the NASA F-111 transonic aircraft technology (TACT) airplane was modified to provide flexible leading and trailing edge flaps; this modified wing is known as the mission adaptive wing (MAW). A dual digital primary fly-by-wire flight control system was developed with analog backup reversion for redundancy. This report discusses the functions, design, and redundancy management of the flight control system for these flaps.

KSC-04pd0458

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility lean toward the body flap to be installed on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0456

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the body flap is moved into position for installation on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0454

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - A Hyster forklift in the Orbiter Processing Facility lifts the body flap to be installed on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0453

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - Workers in the Orbiter Processing Facility help prepare the body flap for lifting prior to installation on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0457

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, the body flap is moved into position for installation on the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0455

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - A Hyster forklift in the Orbiter Processing Facility moves the body flap toward the aft of the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

KSC-04pd0460

NASA Image and Video Library

2004-03-12

KENNEDY SPACE CENTER, FLA. - In the Orbiter Processing Facility, A Hyster forklift supports the body flap as workers secure it to the orbiter Discovery. The body flap is an aluminum structure consisting of ribs, spars, skin panels and a trailing edge assembly. It thermally shields the three main engines during entry and provides pitch control trim during landing approach. Discovery is being processed for launch on the first Return to Flight mission, STS-114.

Airfoil/Wing Flow Control Using Flexible Extended Trailing Edge

DTIC Science & Technology

2009-02-27

and (b) Power spectrums of drag coefficient Figure 4. Mean velocity profiles O Baseline NACA0012. AoA 18 deg c Baseline NACA0012. AoA 20...dynamics, (a) fin amplitude and (b) power spectrum of fin amplitude Development of Computational Tools Simulations of the time-dependent deformation of...combination of experimental, computational and theoretical methods. Compared with Gurney flap and conventional flap, this device enhanced lift at a smaller

Management of Vortices Trailing Flapped Wings via Separation Control

NASA Technical Reports Server (NTRS)

Greenblatt, David

2005-01-01

A pilot study was conducted on a flapped semi-span model to investigate the concept and viability of near-wake vortex management via separation control. Passive control was achieved by means of a simple fairing and active control was achieved via zero mass-flux blowing slots. Vortex sheet strength, estimated by integrating surface pressure ports, was used to predict vortex characteristics by means of inviscid rollup relations. Furthermore, vortices trailing the flaps were mapped using a seven-hole probe. Separation control was found to have a marked effect on vortex location, strength, tangential velocity, axial velocity and size over a wide range of angles of attack and control conditions. In general, the vortex trends were well predicted by the inviscid rollup relations. Manipulation of the separated flow near the flap edges exerted significant control over both outboard and inboard edge vortices while producing negligible lift excursions. Dynamic separation and attachment control was found to be an effective means for dynamically perturbing the vortex from arbitrarily long wavelengths down to wavelengths less than a typical wingspan. In summary, separation control has the potential for application to time-independent or time-dependent wake alleviation schemes, where the latter can be deployed to minimize adverse effects on ride-quality and dynamic structural loading.

Full-scale semi-span tests of an advanced NLF business jet wing

NASA Technical Reports Server (NTRS)

Hahne, David E.; Jordan, Frank L., Jr.; Davis, Patrick J.; Muchmore, C. Byram

1987-01-01

An investigation has been conducted in the NASA Langley Research Center's 30- by 60-Foot Wind Tunnel on a full-scale semispan model to evaluate and document the low-speed, high-lift characteristics of a business-jet class wing utilizing the HSNLF(1)-0213 airfoil section and a single slotted flap system. In addition to the high-lift studies, evaluations of boundary layer transition effects, the effectiveness of a segmented leading-edge droop for improved stall/spin resistance, and roll control effectiveness with and without flap deflection were made. The wind-tunnel investigation showed that deployment of a single-slotted trailing-edge flap provided substantial increments in lift. Fixed transition studies indicated no adverse effects on lift and pitching-moment characteristics for either the cruise or landing configuration. Subscale roll damping tests also indicated that stall/spin resistance could be enhanced through the use of a properly designed leading-edge droop.

Validation of a pair of computer codes for estimation and optimization of subsonic aerodynamic performance of simple hinged-flap systems for thin swept wings

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.

1988-01-01

Extensive correlations of computer code results with experimental data are employed to illustrate the use of linearized theory attached flow methods for the estimation and optimization of the aerodynamic performance of simple hinged flap systems. Use of attached flow methods is based on the premise that high levels of aerodynamic efficiency require a flow that is as nearly attached as circumstances permit. A variety of swept wing configurations are considered ranging from fighters to supersonic transports, all with leading- and trailing-edge flaps for enhancement of subsonic aerodynamic efficiency. The results indicate that linearized theory attached flow computer code methods provide a rational basis for the estimation and optimization of flap system aerodynamic performance at subsonic speeds. The analysis also indicates that vortex flap design is not an opposing approach but is closely related to attached flow design concepts. The successful vortex flap design actually suppresses the formation of detached vortices to produce a small vortex which is restricted almost entirely to the leading edge flap itself.

Airfoil modification effects on subsonic and transonic pressure distributions and performance for the EA-6B airplane

NASA Technical Reports Server (NTRS)

Allison, Dennis O.; Sewall, William G.

1995-01-01

Longitudinal characteristics and wing-section pressure distributions are compared for the EA-6B airplane with and without airfoil modifications. The airfoil modifications were designed to increase low-speed maximum lift for maneuvering, while having a minimal effect on transonic performance. Section contour changes were confined to the leading-edge slat and trailing-edge flap regions of the wing. Experimental data are analyzed from tests in the Langley 16-Foot Transonic Tunnel on the baseline and two modified wing-fuselage configurations with the slats and flaps in their retracted positions. Wing modification effects on subsonic and transonic performance are seen in wing-section pressure distributions of the various configurations at similar lift coefficients. The modified-wing configurations produced maximum lift coefficients which exceeded those of the baseline configuration at low-speed Mach numbers (0.300 and 0.400). This benefit was related to the behavior of the wing upper surface leading-edge suction peak and the behavior of the trailing-edge pressure. At transonic Mach numbers (0.725 to 0.900), the wing modifications produced a somewhat stronger nose-down pitching moment, a slightly higher drag at low-lift levels, and a lower drag at higher lift levels.

Large-Scale Wind-Tunnel Tests of an Airplane Model with an Unswept, Aspect-Ratio-10 Wing, Two Propellers, and Blowing Flaps

NASA Technical Reports Server (NTRS)

Griffin, Roy N., Jr.; Holzhauser, Curt A.; Weiberg, James A.

1958-01-01

An investigation was made to determine the lifting effectiveness and flow requirements of blowing over the trailing-edge flaps and ailerons on a large-scale model of a twin-engine, propeller-driven airplane having a high-aspect-ratio, thick, straight wing. With sufficient blowing jet momentum to prevent flow separation on the flap, the lift increment increased for flap deflections up to 80 deg (the maximum tested). This lift increment also increased with increasing propeller thrust coefficient. The blowing jet momentum coefficient required for attached flow on the flaps was not significantly affected by thrust coefficient, angle of attack, or blowing nozzle height.

Subsonic Investigation of a Leading-Edge Boundary Layer Control Suction System on a High-Speed Civil Transport Configuration

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Applin, Zachary T.; Kemmerly, Guy T.; Coe, Paul L., Jr.; Owens, D. Bruce; Gile, Brenda E.; Parikh, Pradip G.; Smith, Don

1999-01-01

A wind tunnel investigation of a leading edge boundary layer control system was conducted on a High Speed Civil Transport (HSCT) configuration in the Langley 14- by 22-Foot Subsonic Tunnel. Data were obtained over a Mach number range of 0.08 to 0.27, with corresponding chord Reynolds numbers of 1.79 x 10(exp 6) to 5.76 x 10(exp 6). Variations in the amount of suction, as well as the size and location of the suction area, were tested with outboard leading edge flaps deflected 0 and 30 deg and trailing-edge flaps deflected 0 and 20 deg. The longitudinal and lateral aerodynamic data are presented without analysis. A complete tabulated data listing is also presented herein.

Presentation of the acoustic and aerodynamic results of the Aladin 2 concept qualification testing

NASA Technical Reports Server (NTRS)

Collard, M.; Doyotte, C.; Sagner, M.

1985-01-01

Wind tunnel tests were conducted of a scale model of the Aladin 2 aircraft. The propulsion system configuration is described and the air flow caused by jet ejection is analyzed. Three dimensional flow studies in the vicinity of the engine installation were made. Diagrams of the leading and trailing edge flaps are provided. Graphs are developed to show the aerodynamic performance under conditions of various airspeed and flap deflection.

Elastically Shaped Wing Optimization and Aircraft Concept for Improved Cruise Efficiency

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Trinh, Khanh; Reynolds, Kevin; Kless, James; Aftosmis, Michael; Urnes, James, Sr.; Ippolito, Corey

2013-01-01

This paper presents the findings of a study conducted tn 2010 by the NASA Innovation Fund Award project entitled "Elastically Shaped Future Air Vehicle Concept". The study presents three themes in support of meeting national and global aviation challenges of reducing fuel burn for present and future aviation systems. The first theme addresses the drag reduction goal through innovative vehicle configurations via non-planar wing optimization. Two wing candidate concepts have been identified from the wing optimization: a drooped wing shape and an inflected wing shape. The drooped wing shape is a truly biologically inspired wing concept that mimics a seagull wing and could achieve about 5% to 6% drag reduction, which is aerodynamically significant. From a practical perspective, this concept would require new radical changes to the current aircraft development capabilities for new vehicles with futuristic-looking wings such as this concept. The inflected wing concepts could achieve between 3% to 4% drag reduction. While the drag reduction benefit may be less, the inflected-wing concept could have a near-term impact since this concept could be developed within the current aircraft development capabilities. The second theme addresses the drag reduction goal through a new concept of elastic wing shaping control. By aeroelastically tailoring the wing shape with active control to maintain optimal aerodynamics, a significant drag reduction benefit could be realized. A significant reduction in fuel burn for long-range cruise from elastic wing shaping control could be realized. To realize the potential of the elastic wing shaping control concept, the third theme emerges that addresses the drag reduction goal through a new aerodynamic control effector called a variable camber continuous trailing edge flap. Conventional aerodynamic control surfaces are discrete independent surfaces that cause geometric discontinuities at the trailing edge region. These discontinuities promote vorticities which result in drag rises as well as noise sources. The variable camber trailing edge flap concept could provide a substantial drag reduction benefit over a conventional discrete flap system. Aerodynamic simulations show a drag reduction of over 50% could be achieved with the flap concept over a conventional discrete flap system.

Vortical flow management for improved configuration aerodynamics: Recent experiences

NASA Technical Reports Server (NTRS)

Rao, D. M.

1983-01-01

Recent progress in vortex-control applications for alleviating the adverse consequences of three dimensional separation and vortical interactions on slender body/swept wing configurations is reported. Examples include helical separation trip to alleviate the side force due to forebody vortex asymmetry; hinged strakes to avoid vortex breakdown effects; compartmentation of swept leading edge separation to delay the pitch-up instability; under wing vortex trip and vortex trip and vortex flaps for drag reduction at high lift; and an apex-flap trimmer to fully utilize the lift capability of trailing-edge flaps for take off and landing of delta wings. Experimental results on generic wind-tunnel models are presented to illustrate the vortex-management concepts involved and to indicate their potential for enhancing the subsonic aerodynamics of supersonic-cruise type vehicles.

Pressure distribution data from tests of 2.29-meter (7.5-ft.) span EET high-lift research model in Langley 4- by 7-meter tunnel

NASA Technical Reports Server (NTRS)

Morgan, H. L., Jr.

1982-01-01

A 2.29 m (7.5 ft.) span high-lift research model equipped with full-span leading-edge slat and part-span double-slotted trailing-edge flap was tested in the Langley 4- by 7-Meter Tunnel to determine the low speed performance characteristics of a representative high aspect ratio suprcritical wing. These tests were performed in support of the Energy Efficient Transport (EET) program which is one element of the Aircraft Energy Efficiency (ACEE) project. Static longitudinal forces and moments and chordwise pressure distributions at three spanwise stations were measured for cruise, climb, two take-off flap, and two landing flap wing configurations. The tabulated and plotted pressure distribution data is presented without analysis or discussion.

Pressure distribution data from tests of 2.29 M (7.5 feet) span EET high-lift transport aircraft model in the Ames 12-foot pressure tunnel

NASA Technical Reports Server (NTRS)

Kjelgaard, S. O.; Morgan, H. L., Jr.

1983-01-01

A high-lift transport aircraft model equipped with full-span leading-edge slat and part-span double-slotted trailing-edge flap was tested in the Ames 12-ft pressure tunnel to determine the low-speed performance characteristics of a representative high-aspect-ratio supercritical wing. These tests were performed in support of the Energy Efficient Transport (EET) program which is one element of the Aircraft Energy Efficiency (ACEE) project. Static longitudinal forces and moments and chordwise pressure distributions at three spanwise stations were measured for cruise, climb, two take-off flap, and two landing flap wing configurations. The tabulated and plotted pressure distribution data is presented without analysis or discussion.

VIEW OF LEFT WING AND FUSELAGE FROM TOP LEVEL OF ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

VIEW OF LEFT WING AND FUSELAGE FROM TOP LEVEL OF TAIL DOCK STAND. LEADING AND TRAILING EDGE FLAPS ARE DOWN; AIELERONS ARE IN NEUTRAL. ENGINE COWLING OFF FOR HEAVY INSPECTION. - Greater Buffalo International Airport, Maintenance Hangar, Buffalo, Erie County, NY

Lift-Enhancing Tabs on Multielement Airfoils

NASA Technical Reports Server (NTRS)

Ross, James C.; Storms, Bruce L.; Carrannanto, Paul G.

1995-01-01

The use of flat-plate tabs (similar to Gurney flaps) to enhance the lift of multielement airfoils is extended here by placing them on the pressure side and near the trailing edge of the main element rather than just on the furthest downstream wing element. The tabs studied range in height from 0.125 to 1.25% of the airfoil reference chord. In practice, such tabs would be retracted when the high-lift system is stowed. The effectiveness of the concept was demonstrated experimentally and computationally on a two-dimensional NACA 63(sub 2)-215 Mod B airfoil with a single-slotted, 30%-chord flap. Both the experiments and computations showed that the tabs significantly increase the lift at a given angle of attack and the maximum lift coefficient of the airfoil. The computational results showed that the increased lift was a result of additional turning of the flow by the tab that reduced or eliminated now separation on the flap. The best configuration tested, a 0.5%-chord tab placed 0.5% chord upstream of the trailing edge of the main element, increased the maximum lift coefficient of the airfoil by 12% and the maximum lift-to-drag ratio by 40%.

Aeroelastic Airworthiness Assesment of the Adaptive Compliant Trailing Edge Flaps

NASA Technical Reports Server (NTRS)

Herrera, Claudia Y.; Spivey, Natalie D.; Lung, Shun-fat; Ervin, Gregory; Flick, Peter

2015-01-01

The Adaptive Compliant Trailing Edge (ACTE) demonstrator is a joint task under the National Aeronautics and Space Administration Environmentally Responsible Aviation Project in partnership with the Air Force Research Laboratory and FlexSys, Inc. (Ann Arbor, Michigan). The project goal is to develop advanced technologies that enable environmentally friendly aircraft, such as adaptive compliant technologies. The ACTE demonstrator flight-test program encompassed replacing the Fowler flaps on the SubsoniC Aircraft Testbed, a modified Gulfstream III (Gulfstream Aerospace, Savannah, Georgia) aircraft, with control surfaces developed by FlexSys. The control surfaces developed by FlexSys are a pair of uniquely-designed unconventional flaps to be used as lifting surfaces during flight-testing to validate their structural effectiveness. The unconventional flaps required a multidisciplinary airworthiness assessment to prove they could withstand the prescribed flight envelope. Several challenges were posed due to the large deflections experienced by the structure, requiring non-linear analysis methods. The aeroelastic assessment necessitated both conventional and extensive testing and analysis methods. A series of ground vibration tests (GVTs) were conducted to provide modal characteristics to validate and update finite element models (FEMs) used for the flutter analyses for a subset of the various flight configurations. Numerous FEMs were developed using data from FlexSys and the ground tests. The flap FEMs were then attached to the aircraft model to generate a combined FEM that could be analyzed for aeroelastic instabilities. The aeroelastic analysis results showed the combined system of aircraft and flaps were predicted to have the required flutter margin to successfully demonstrate the adaptive compliant technology. This paper documents the details of the aeroelastic airworthiness assessment described, including the ground testing and analyses, and subsequent flight-testing performed on the unconventional ACTE flaps.

SMART Rotor Development and Wind-Tunnel Test

NASA Technical Reports Server (NTRS)

Lau, Benton H.; Straub, Friedrich; Anand, V. R.; Birchette, Terry

2009-01-01

Boeing and a team from Air Force, NASA, Army, Massachusetts Institute of Technology, University of California at Los Angeles, and University of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center, figure 1. The SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing-edge flap on each blade. The development effort included design, fabrication, and component testing of the rotor blades, the trailing-edge flaps, the piezoelectric actuators, the switching power amplifiers, the actuator control system, and the data/power system. Development of the smart rotor culminated in a whirl-tower hover test which demonstrated the functionality, robustness, and required authority of the active flap system. The eleven-week wind tunnel test program evaluated the forward flight characteristics of the active-flap rotor, gathered data to validate state-of-the-art codes for rotor noise analysis, and quantified the effects of open- and closed-loop active-flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness and the reliability of the flap actuation system were successfully demonstrated in more than 60 hours of wind-tunnel testing. The data acquired and lessons learned will be instrumental in maturing this technology and transitioning it into production. The development effort, test hardware, wind-tunnel test program, and test results will be presented in the full paper.

Full-scale wind-tunnel tests of high-lift system modifications on a carrier based fighter aircraft

NASA Technical Reports Server (NTRS)

Meyn, Larry A.; Zell, Peter T.; Hagan, John L.; Schoch, David

1993-01-01

Modifications to the high-lift system of a full-scale F/A-I8A were tested in the 80- by 120-Foot Wind Tunnel of the National Full-Scale Aerodynamics Complex at the NASA Ames Research Center in Moffett Field, California. The objective was to measure the effect of simple modifications on the aerodynamic performance of the high-lift system. The modifications included the placement of a straight fairing in the shroud cove above the trailing-edge flap and the addition of seals to prevent air leakage through the hinge lines of the leading-edge flap, the trailing-edge shroud, and the wing fold. The test was carried out on an actual F/A-18A with it's flaps deployed in the landing approach configuration. The angle of attack ranged from 0 to 16 degrees and the wind speed was 100 knots. At an angle of attack of 8 degrees, the trimmed lift coefficient was improved by 0.09 with all wing seals in place. This corresponds to a reduction in the approach speed for the F/A-I8A of about 5 knots. The seal along the wing fold hinge, a feature present on many naval aircraft, provided one third of the total increment in trimmed lift. A comparison of the full-scale wind-tunnel results with those obtained from flight test is also presented.

Experimental and numerical studies of beetle-inspired flapping wing in hovering flight.

PubMed

Van Truong, Tien; Le, Tuyen Quang; Park, Hoon Cheol; Byun, Doyoung

2017-05-17

In this paper, we measure unsteady forces and visualize 3D vortices around a beetle-like flapping wing model in hovering flight by experiment and numerical simulation. The measurement of unsteady forces and flow patterns around the wing were conducted using a dynamically scaled wing model in the mineral-oil tank. The wing kinematics were directly derived from the experiment of a real beetle. The 3D flow structures of the flapping wing were captured by using air bubble visualization while forces were measured by a sensor attached at the wing base. In comparison, the size and topology of spiral leading edge vortex, trailing edge vortex and tip vortex are well matched from experimental and numerical studies. In addition, the time history of forces calculated from numerical simulation is also similar to that from theforce measurement. A difference of average force is in order of 10 percent. The results indicate that the leading edge vortex due to rotational acceleration at the end of the stroke during flapping wing causes significant reduction of lift. The present study provides useful information on hover flight to develop a beetle-like flapping wing Micro Air Vehicle.

An Investigation of the Wing and the Wing-Fuselage Combination of a Full-Scale Model of the Republic XP-91 Airplane in the Ames 40-by 80-Foot Wing Tunnel

NASA Technical Reports Server (NTRS)

Hunton, Lynn W.; Dew, Joseph K.

1948-01-01

Wind-tunnel tests of a full-scale model of the Republic XP-91 airplane were conducted to determine the longitudinal and lateral characteristics of the wing alone and the wing-fuselage combination, the characteristics of the aileron, and the damping in roll af the wing alone. Various high-lift devices were investigated including trailing-edge split flaps and partial- and full-span leading-edge slats and Krueger-type nose flaps. Results of this investigation showed that a very significant gain in maximum lift could be achieved through use of the proper leading-edge device, The maximum lift coefficient of the model with split flaps and the original partial-span straight slats was only 1.2; whereas a value of approximately 1.8 was obtained by drooping the slat and extending it full span, Improvement in maximum lift of approximately the same amount resulted when a full-span nose flap was substituted for the original partial-span slat.

Aeroelastic Analysis of a Flexible Wing Wind Tunnel Model with Variable Camber Continuous Trailing Edge Flap Design

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Lebofsky, Sonia

2015-01-01

This paper presents data analysis of a flexible wing wind tunnel model with a variable camber continuous trailing edge flap (VCCTEF) design for drag minimization tested at the University of Washington Aeronautical Laboratory (UWAL). The wind tunnel test was designed to explore the relative merit of the VCCTEF concept for improved cruise efficiency through the use of low-cost aeroelastic model test techniques. The flexible wing model is a 10%-scale model of a typical transport wing and is constructed of woven fabric composites and foam core. The wing structural stiffness in bending is tailored to be half of the stiffness of a Boeing 757-era transport wing while the torsional stiffness is about the same. This stiffness reduction results in a wing tip deflection of about 10% of the wing semi-span. The VCCTEF is a multi-segment flap design having three chordwise camber segments and five spanwise flap sections for a total of 15 individual flap elements. The three chordwise camber segments can be positioned appropriately to create a desired trailing edge camber. Elastomeric material is used to cover the gaps in between the spanwise flap sections, thereby creating a continuous trailing edge. Wind tunnel data analysis conducted previously shows that the VCCTEF can achieve a drag reduction of up to 6.31% and an improvement in the lift-to-drag ratio (L=D) of up to 4.85%. A method for estimating the bending and torsional stiffnesses of the flexible wingUWAL wind tunnel model from static load test data is presented. The resulting estimation indicates that the stiffness of the flexible wing is significantly stiffer in torsion than in bending by as much as 9 to 1. The lift prediction for the flexible wing is computed by a coupled aerodynamic-structural model. The coupled model is developed by coupling a conceptual aerodynamic tool Vorlax with a finite-element model of the flexible wing via an automated geometry deformation tool. Based on the comparison of the lift curve slope, the lift prediction for the rigid wing is in good agreement with the estimated lift coefficients derived from the wind tunnel test data. Due to the movement of the VCCTEF during the wind tunnel test, uncertainty in the lift prediction due to the indicated variations of the VCCTEF deflection is studied. The results show a significant spread in the lift prediction which contradicts the consistency in the aerodynamic measurements, thus suggesting that the indicated variations as measured by the VICON system may not be reliable. The lift prediction of the flexible wing agrees very well with the measured lift curve for the baseline configuration. The computed bending deflection and wash-out twist of the flexible wing also match reasonably well with the aeroelastic deflection measurements. The results demonstrate the validity of the aerodynamic-structural tool for use to analyze aerodynamic performance of flexible wings.

STOL Characteristics of a Propeller-Driven, Aspect-Ratio-10, Straight-Wing Airplane with Boundary-Layer Control Flaps, as Estimated from Large-Scale Wind-Tunnel Tests

NASA Technical Reports Server (NTRS)

Weiberg, James A; Holzhauser, Curt A.

1961-01-01

A study is presented of the improvements in take-off and landing distances possible with a conventional propeller-driven transport-type airplane when the available lift is increased by propeller slipstream effects and by very effective trailing-edge flaps and ailerons. This study is based on wind-tunnel tests of a 45-foot span, powered model, with BLC on the trailing-edge flaps and controls. The data were applied to an assumed airplane with four propellers and a wing loading of 50 pounds per square foot. Also included is an examination of the stability and control problems that may result in the landing and take-off speed range of such a vehicle. The results indicated that the landing and take-off distances could be more than halved by the use of highly effective flaps in combination with large amounts of engine power to augment lift (STOL). At the lowest speeds considered (about 50 knots), adequate longitudinal stability was obtained but the lateral and directional stability were unsatisfactory. At these low speeds, the conventional aerodynamic control surfaces may not be able to cope with the forces and moments produced by symmetric, as well as asymmetric, engine operation. This problem was alleviated by BLC applied to the control surfaces.

Flexible wings in flapping flight

NASA Astrophysics Data System (ADS)

Moret, Lionel; Thiria, Benjamin; Zhang, Jun

2007-11-01

We study the effect of passive pitching and flexible deflection of wings on the forward flapping flight. The wings are flapped vertically in water and are allowed to move freely horizontally. The forward speed is chosen by the flapping wing itself by balance of drag and thrust. We show, that by allowing the wing to passively pitch or by adding a flexible extension at its trailing edge, the forward speed is significantly increased. Detailed measurements of wing deflection and passive pitching, together with flow visualization, are used to explain our observations. The advantage of having a wing with finite rigidity/flexibility is discussed as we compare the current results with our biological inspirations such as birds and fish.

McDonnell F4H Model in Ames 40X80 foot Wind Tunnel.

NASA Image and Video Library

1956-10-19

Application of blowing type boundry-layer control to the leading and trailing edge flaps of a 52 deg swept wing. 3/4 view of Aspect Ratio 2.8, taper ratio .17, 45 deg swept back wing model -3/4 front view

Aerodynamic Flight-Test Results for the Adaptive Compliant Trailing Edge

NASA Technical Reports Server (NTRS)

Cumming, Stephen B.; Smith, Mark S.; Ali, Aliyah N.; Bui, Trong T.; Ellsworth, Joel C.; Garcia, Christian A.

2016-01-01

The aerodynamic effects of compliant flaps installed onto a modified Gulfstream III airplane were investigated. Analyses were performed prior to flight to predict the aerodynamic effects of the flap installation. Flight tests were conducted to gather both structural and aerodynamic data. The airplane was instrumented to collect vehicle aerodynamic data and wing pressure data. A leading-edge stagnation detection system was also installed. The data from these flights were analyzed and compared with predictions. The predictive tools compared well with flight data for small flap deflections, but differences between predictions and flight estimates were greater at larger deflections. This paper describes the methods used to examine the aerodynamics data from the flight tests and provides a discussion of the flight-test results in the areas of vehicle aerodynamics, wing sectional pressure coefficient profiles, and air data.

Low-speed wind-tunnel investigation of the stability and control characteristics of a series of flying wings with sweep angles of 50 deg

NASA Technical Reports Server (NTRS)

Fears, Scott P.; Ross, Holly M.; Moul, Thomas M.

1995-01-01

A wind-tunnel investigation was conducted in the Langley 12-Foot Low-Speed Tunnel to study the low-speed stability and control characteristics of a series of four flying wings over an extended range of angle of attack (-8 deg to 48 deg). Because of the current emphasis on reducing the radar cross section (RCS) of new military aircraft, the planform of each wing was composed of lines swept at a relatively high angle of 50 deg, and all the trailing-edge lines were aligned with one of the two leading edges. Three arrow planforms with different aspect ratios and one diamond planform were tested. The models incorporated leading-edge flaps for improved longitudinal characteristics and lateral stability and had trailing-edge flaps in three segments that were deflected differentially for roll control, symmetrically for pitch control, and in a split fashion for yaw control. Three top body widths and two sizes of twin vertical tails were also tested on each model. A large aerodynamic database was compiled that could be used to evaluate some of the trade-offs involved in the design of a configuration with a reduced RCS and good flight dynamic characteristics.

Low-speed wind-tunnel investigation of the stability and control characteristics of a series of flying wings with sweep angles of 70 deg

NASA Technical Reports Server (NTRS)

Ross, Holly M.; Fears, Scott P.; Moul, Thomas M.

1995-01-01

A wind-tunnel investigation was conducted in the Langley 12-Foot Low-Speed Tunnel to study the low-speed stability and control characteristics of a series of four flying wings over an extended range of angle of attack (-8 deg to 48 deg). Because of the current emphasis on reducing the radar cross section (RCS) of new military aircraft, the planform of each wing was composed of lines swept at a relatively high angle of 70 deg, and all the trailing edges and control surface hinge lines were aligned with one of the two leading edges. Three arrow planforms with different aspect ratios and one diamond planform were tested. The models incorporated leading-edge flaps for improved longitudinal characteristics and lateral stability and had three sets of trailing-edge flaps that were deflected differentially for roll control, symmetrically for pitch control, and in a split fashion for yaw control. Three top body widths and two sizes of twin vertical tails were also tested on each model. A large aerodynamic database was compiled that could be used to evaluate some of the trade-offs involved in the design of a configuration with a reduced RCS and good flight dynamic characteristics.

Low-speed wind tunnel investigation of the stability and control characteristics of a series of flying wings with sweep angles of 60 deg

NASA Technical Reports Server (NTRS)

Moul, Thomas M.; Fears, Scott P.; Ross, Holly M.; Foster, John V.

1995-01-01

A wind tunnel investigation was conducted in the Langley 12-Foot Low-Speed Wind Tunnel to study the low-speed stability and control characteristics of a series of four flying wings over an extended range of angle of attack (-8 deg to 48 deg). Because of the current emphasis on reducing the radar cross section of new military aircraft, the planform of each wing was composed of lines swept at a relatively high angle of 60 deg, and all the trailing-edge lines were aligned with one of the two leading edges. Three arrow planforms with different aspect ratios and one diamond planform were tested. The models incorporated leading-edge flaps for improved pitching-moment characteristics and lateral stability and had three sets of trailing-edge flaps that were deflected differentially for roll control, symmetrically for pitch control, and in a split fashion for yaw control. Top bodies of three widths and twin vertical tails of various sizes and locations were also tested on each model. A large aerodynamic database was compiled that could be used to evaluate some of the trade-offs involved in the design of a configuration with a reduced radar cross section and good flight dynamic characteristics.

ARC-1961-A-28387

NASA Image and Video Library

1961-10-31

Lockheed NC-130B STOL turboprop-powered aircraft with ailerons drooped 30 degrees. Note trailing-edge flaps deflected 90 degrees for increased lift. Two T-56 turboshaft engines, which drove wing-mounted load compressors for boundary-layer control, are mounted on outboard wing pods. Landing approach speed was reduced 30 knots with boundary-layer control

Analysis of a Multi-Flap Control System for a Swashplateless Rotor

NASA Technical Reports Server (NTRS)

Sekula, Martin K.; Wilbur, Matthew L.

2011-01-01

An analytical study was conducted examining the feasibility of a swashplateless rotor controlled through two trailing edge flaps (TEF), where the cyclic and collective controls were provided by separate TEFs. This analysis included a parametric study examining the impact of various design parameters on TEF deflections. Blade pitch bearing stiffness; blade pitch index; and flap chord, span, location, and control function of the inboard and outboard flaps were systematically varied on a utility-class rotorcraft trimmed in steady level flight. Gradient-based optimizations minimizing flap deflections were performed to identify single- and two-TEF swashplateless rotor designs. Steady, forward and turning flight analyses suggest that a two-TEF swashplateless rotor where the outboard flap provides cyclic control and inboard flap provides collective control can reduce TEF deflection requirements without a significant impact on power, compared to a single-TEF swashplateless rotor design.

A test of a vortex method for the computation of flap side edge noise

NASA Technical Reports Server (NTRS)

Martin, James E.

1995-01-01

Upon approach to landing, a major source location of airframe noise occurs at the side edges of the part span, trailing edge flaps. In the vicinity of these flaps, a complex arrangement of spanwise flow with primary and secondary tip vortices may form. Each of these vortices is observed to become fully three-dimensional. In the present study, a numerical model is developed to investigate the noise radiated from the side edge of a flap. The inherent three-dimensionality of this flow forces us to carefully consider a numerical scheme which will be both accurate in its prediction of the flow acoustics and also computationally efficient. Vortex methods have offered a fast and efficient means of simulating many two and three-dimensional, vortex dominated flows. In vortex methods, the time development of the flow is tracked by following exclusively the vorticity containing regions. Through the Biot-Savart law, knowledge of the vorticity field enables one to obtain flow quantities at any desired location during the flow evolution. In the present study, a numerical procedure has been developed which incorporates the Lagrangian approach of vortex methods into a calculation for the noise radiated by a flow-surface interaction. In particular, the noise generated by a vortex in the presence of a flat half plane is considered. This problem serves as a basic model of flap edge flow. It also permits the direct comparison between our computed results and previous acoustic analyses performed for this problem. In our numerical simulations, the mean flow is represented by the complex potential W(z) = Aiz(exp l/2), which is obtained through conformal mapping techniques. The magnitude of the mean flow is controlled by the parameter A. This mean flow has been used in the acoustic analysis by Hardin and is considered a reasonable model of the flow field in the vicinity of the edge and away from the leading and trailing edges of the flap. To represent the primary vortex which occurs near the flap, a point vortex is introduced just below the flat half plane. Using a technique from panel methods, boundary conditions on the flap surface are satisfied by the introduction of a row of stationary point vortices along the extent of the flap. At each time step in the calculation, the strength of these vortices is chosen to eliminate the normal velocity at intermediary collocation points. The time development of the overall flow field is then tracked using standard techniques from vortex methods. Vortex trajectories obtained through this computation are in good agreement with those predicted by the analytical solution given by Hardin, thus verifying the viability of this procedure for more complex flow arrangements. For the flow acoustics, the Ffowcs Williams-Hawkings equation is numerically integrated. This equation supplies the far field acoustic pressure based upon pressures occurring along the flap surface. With our vortex method solution, surface pressures may be obtained with exceptional resolution. The Ffowcs Williams-Hawkings equation is integrated using a spatially fourth order accurate Simpson's rule. Rational function interpolation is used to obtain the surface pressures at the appropriate retarded times. Comparisons between our numerical results for the acoustic pressure and those predicted by the Hardin analysis have been made. Preliminary results indicate the need for an improved integration technique. In the future, the numerical procedure developed in this study will be applied to the case of a rectangular flap of finite thickness and ultimately modified for application to the fully three-dimensional problem.

Static Internal Performance of a Two-Dimensional Convergent-Divergent Nozzle with External Shelf

NASA Technical Reports Server (NTRS)

Lamb, Milton; Taylor, John G.; Frassinelli, Mark C.

1996-01-01

An investigation was conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance of a two-dimensional convergent-divergent nozzle. The nozzle design was tested with dry and afterburning throat areas, which represent different power settings and three expansion ratios. For each of these configurations, three trailing-edge geometries were tested. The baseline geometry had a straight trailing edge. Two different shaping techniques were applied to the baseline nozzle design to reduce radar observables: the scarfed design and the sawtooth design. A flat plate extended downstream of the lower divergent flap trailing edge parallel to the model centerline to form a shelf-like expansion surface. This shelf was designed to shield the plume from ground observation (infrared radiation (IR) signature suppression). The shelf represents the part of the aircraft structure that might be present in an installed configuration. These configurations were tested at nozzle pressure ratios from 2.0 to 12.0.

Incremental wind tunnel testing of high lift systems

NASA Astrophysics Data System (ADS)

Victor, Pricop Mihai; Mircea, Boscoianu; Daniel-Eugeniu, Crunteanu

2016-06-01

Efficiency of trailing edge high lift systems is essential for long range future transport aircrafts evolving in the direction of laminar wings, because they have to compensate for the low performance of the leading edge devices. Modern high lift systems are subject of high performance requirements and constrained to simple actuation, combined with a reduced number of aerodynamic elements. Passive or active flow control is thus required for the performance enhancement. An experimental investigation of reduced kinematics flap combined with passive flow control took place in a low speed wind tunnel. The most important features of the experimental setup are the relatively large size, corresponding to a Reynolds number of about 2 Million, the sweep angle of 30 degrees corresponding to long range airliners with high sweep angle wings and the large number of flap settings and mechanical vortex generators. The model description, flap settings, methodology and results are presented.

Refined AFC-Enabled High-Lift System Integration Study

NASA Technical Reports Server (NTRS)

Hartwich, Peter M.; Shmilovich, Arvin; Lacy, Douglas S.; Dickey, Eric D.; Scalafani, Anthony J.; Sundaram, P.; Yadlin, Yoram

2016-01-01

A prior trade study established the effectiveness of using Active Flow Control (AFC) for reducing the mechanical complexities associated with a modern high-lift system without sacrificing aerodynamic performance at low-speed flight conditions representative of takeoff and landing. The current technical report expands on this prior work in two ways: (1) a refined conventional high-lift system based on the NASA Common Research Model (CRM) is presented that is more representative of modern commercial transport aircraft in terms of stall characteristics and maximum Lift/Drag (L/D) ratios at takeoff and landing-approach flight conditions; and (2) the design trade space for AFC-enabled high-lift systems is expanded to explore a wider range of options for improving their efficiency. The refined conventional high-lift CRM (HL-CRM) concept features leading edge slats and slotted trailing edge flaps with Fowler motion. For the current AFC-enhanced high lift system trade study, the refined conventional high-lift system is simplified by substituting simply-hinged trailing edge flaps for the slotted single-element flaps with Fowler motion. The high-lift performance of these two high-lift CRM variants is established using Computational Fluid Dynamics (CFD) solutions to the Reynolds-Averaged Navier-Stokes (RANS) equations. These CFD assessments identify the high-lift performance that needs to be recovered through AFC to have the CRM variant with the lighter and mechanically simpler high-lift system match the performance of the conventional high-lift system. In parallel to the conventional high-lift concept development, parametric studies using CFD guided the development of an effective and efficient AFC-enabled simplified high-lift system. This included parametric trailing edge flap geometry studies addressing the effects of flap chord length and flap deflection. As for the AFC implementation, scaling effects (i.e., wind-tunnel versus full-scale flight conditions) are addressed, as are AFC architecture aspects such as AFC unit placement, number AFC units, operating pressures, mass flow rates, and steady versus unsteady AFC applications. These efforts led to the development of a novel traversing AFC actuation concept which is efficient in that it reduces the AFC mass flow requirements by as much as an order of magnitude compared to previous AFC technologies, and it is predicted to be effective in driving the aerodynamic performance of a mechanical simplified high-lift system close to that of the reference conventional high-lift system. Conceptual system integration studies were conducted for the AFC-enhanced high-lift concept applied to a NASA Environmentally Responsible Aircraft (ERA) reference configuration, the so-called ERA-0003 concept. The results from these design integration assessments identify overall system performance improvement opportunities over conventional high-lift systems that suggest the viability of further technology maturation efforts for AFC-enabled high lift flap systems. To that end, technical challenges are identified associated with the application of AFC-enabled high-lift systems to modern transonic commercial transports for future technology maturation efforts.

The PELskin project-part V: towards the control of the flow around aerofoils at high angle of attack using a self-activated deployable flap.

PubMed

Rosti, Marco E; Kamps, Laura; Bruecker, Christoph; Omidyeganeh, Mohammad; Pinelli, Alfredo

2017-01-01

During the flight of birds, it is often possible to notice that some of the primaries and covert feathers on the upper side of the wing pop-up under critical flight conditions, such as the landing approach or when stalking their prey (see Fig. 1) . It is often conjectured that the feathers pop up plays an aerodynamic role by limiting the spread of flow separation . A combined experimental and numerical study was conducted to shed some light on the physical mechanism determining the feathers self actuation and their effective role in controlling the flow field in nominally stalled conditions. In particular, we have considered a NACA0020 aerofoil, equipped with a flexible flap at low chord Reynolds numbers. A parametric study has been conducted on the effects of the length, natural frequency, and position of the flap. A configuration with a single flap hinged on the suction side at 70 % of the chord size c (from the leading edge), with a length of [Formula: see text] matching the shedding frequency of vortices at stall condition has been found to be optimum in delivering maximum aerodynamic efficiency and lift gains. Flow evolution both during a ramp-up motion (incidence angle from [Formula: see text] to [Formula: see text] with a reduced frequency of [Formula: see text], [Formula: see text] being the free stream velocity magnitude), and at a static stalled condition ([Formula: see text]) were analysed with and without the flap. A significant increase of the mean lift after a ramp-up manoeuvre is observed in presence of the flap. Stall dynamics (i.e., lift overshoot and oscillations) are altered and the simulations reveal a periodic re-generation cycle composed of a leading edge vortex that lift the flap during his passage, and an ejection generated by the relaxing of the flap in its equilibrium position. The flap movement in turns avoid the interaction between leading and trailing edge vortices when lift up and push the trailing edge vortex downstream when relaxing back. This cyclic behaviour is clearly shown by the periodic variation of the lift about the average value, and also from the periodic motion of the flap. A comparison with the experiments shows a similar but somewhat higher non-dimensional frequency of the flap oscillation. By assuming that the cycle frequency scales inversely with the boundary layer thickness, one can explain the higher frequencies observed in the experiments which were run at a Reynolds number about one order of magnitude higher than in the simulations. In addition, in experiments the periodic re-generation cycle decays after 3-4 periods ultimately leading to the full stall of the aerofoil. In contrast, the 2D simulations show that the cycle can become self-sustained without any decay when the flap parameters are accurately tuned.

76 FR 9498 - Airworthiness Directives; The Boeing Company Model 737-300, -400, and -500 Series Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2011-02-18

... installed, repetitive inspections for corrosion and indications of corrosion on affected carriage spindles... repetitive inspections. This AD results from reports of corrosion found on carriage spindles that are located on the outboard trailing edge flaps. We are issuing this AD to detect and correct corrosion of the...

Overview: Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Hashemi, Kelley

2017-01-01

An overview of recent aeroelasitc wing-shaping work at the NASA Ames Research Center is presented. The highlight focuses on activity related to the Performance Adaptive Aeroelastic Wing concept and related Variable Camber Continuous Trailing Edge Flap actuation system. Topics covered include drag-reducing configurations and online algorithms, gust and maneuver load techniques, and wind tunnel demonstrations.

Strain Gage Load Calibration of the Wing Interface Fittings for the Adaptive Compliant Trailing Edge Flap Flight Test

NASA Technical Reports Server (NTRS)

Miller, Eric J.; Holguin, Andrew C.; Cruz, Josue; Lokos, William A.

2014-01-01

The safety-of-flight parameters for the Adaptive Compliant Trailing Edge (ACTE) flap experiment require that flap-to-wing interface loads be sensed and monitored in real time to ensure that the structural load limits of the wing are not exceeded. This paper discusses the strain gage load calibration testing and load equation derivation methodology for the ACTE interface fittings. Both the left and right wing flap interfaces were monitored; each contained four uniquely designed and instrumented flap interface fittings. The interface hardware design and instrumentation layout are discussed. Twenty-one applied test load cases were developed using the predicted in-flight loads. Pre-test predictions of strain gage responses were produced using finite element method models of the interface fittings. Predicted and measured test strains are presented. A load testing rig and three hydraulic jacks were used to apply combinations of shear, bending, and axial loads to the interface fittings. Hardware deflections under load were measured using photogrammetry and transducers. Due to deflections in the interface fitting hardware and test rig, finite element model techniques were used to calculate the reaction loads throughout the applied load range, taking into account the elastically-deformed geometry. The primary load equations were selected based on multiple calibration metrics. An independent set of validation cases was used to validate each derived equation. The 2-sigma residual errors for the shear loads were less than eight percent of the full-scale calibration load; the 2-sigma residual errors for the bending moment loads were less than three percent of the full-scale calibration load. The derived load equations for shear, bending, and axial loads are presented, with the calculated errors for both the calibration cases and the independent validation load cases.

Experimental and Computational Investigation of Lift-Enhancing Tabs on a Multi-Element Airfoil

NASA Technical Reports Server (NTRS)

Ashby, Dale L.

1996-01-01

An experimental and computational investigation of the effect of lift-enhancing tabs on a two-element airfoil has been conducted. The objective of the study was to develop an understanding of the flow physics associated with lift-enhancing tabs on a multi-element airfoil. An NACA 63(2)-215 ModB airfoil with a 30% chord fowler flap was tested in the NASA Ames 7- by 10-Foot Wind Tunnel. Lift-enhancing tabs of various heights were tested on both the main element and the flap for a variety of flap riggings. A combination of tabs located at the main element and flap trailing edges increased the airfoil lift coefficient by 11% relative to the highest lift coefficient achieved by any baseline configuration at an angle of attack of 0 deg, and C(sub 1max) was increased by 3%. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computed results predicted all of the trends observed in the experimental data quite well. In addition, a simple analytic model based on potential flow was developed to provide a more detailed understanding of how lift-enhancing tabs work. The tabs were modeled by a point vortex at the air-foil or flap trailing edge. Sensitivity relationships were derived which provide a mathematical basis for explaining the effects of lift-enhancing tabs on a multi-element airfoil. Results of the modeling effort indicate that the dominant effects of the tabs on the pressure distribution of each element of the airfoil can be captured with a potential flow model for cases with no flow separation.

A Comparative Study of Simulated and Measured Gear-Flap Flow Interaction

NASA Technical Reports Server (NTRS)

Khorrami, Mehdi R.; Mineck, Raymond E.; Yao, Chungsheng; Jenkins, Luther N.; Fares, Ehab

2015-01-01

The ability of two CFD solvers to accurately characterize the transient, complex, interacting flowfield asso-ciated with a realistic gear-flap configuration is assessed via comparison of simulated flow with experimental measurements. The simulated results, obtained with NASA's FUN3D and Exa's PowerFLOW® for a high-fidelity, 18% scale semi-span model of a Gulfstream aircraft in landing configuration (39 deg flap deflection, main landing gear on and off) are compared to two-dimensional and stereo particle image velocimetry measurements taken within the gear-flap flow interaction region during wind tunnel tests of the model. As part of the bench-marking process, direct comparisons of the mean and fluctuating velocity fields are presented in the form of planar contour plots and extracted line profiles at measurement planes in various orientations stationed in the main gear wake. The measurement planes in the vicinity of the flap side edge and downstream of the flap trailing edge are used to highlight the effects of gear presence on tip vortex development and the ability of the computational tools to accurately capture such effects. The present study indicates that both computed datasets contain enough detail to construct a relatively accurate depiction of gear-flap flow interaction. Such a finding increases confidence in using the simulated volumetric flow solutions to examine the behavior of pertinent aer-odynamic mechanisms within the gear-flap interaction zone.

Wind-tunnel tests of a Clark Y wing with 'Maxwell' leading-edge slots

NASA Technical Reports Server (NTRS)

Gauvain, William E

1937-01-01

Aerodynamic force tests of a Clark Y wing equipped with "Maxwell" type leading-edge slots were conducted in the N.A.C.A. 7- by 10-foot tunnel to ascertain the aerodynamic characteristics, which involved the determination of the best slot-gap opening, the effects of slat width, and the effect of a trailing-edge flap. The Maxwell wing with a wide-chord slat (0.30 c(sub w)) and with a 0.211 c(sub w) split flap deflected 60 degrees had a C(sub L sub max) of 2.53 or about twice that of the plain wing. The wing with the wide slat also had, in general, improved aerodynamic characteristics over those of the Maxwell wing with slat, and had about the same aerodynamic characteristics as a Handley Page slotted wing with approximately the same size of slat.

Full-scale semispan tests of a business-jet wing with a natural laminar flow airfoil

NASA Technical Reports Server (NTRS)

Hahne, David E.; Jordan, Frank L., Jr.

1991-01-01

A full-scale semispan model was investigated to evaluate and document the low-speed, high-lift characteristics of a business-jet class wing that utilized the HSNLF(1)-0213 airfoil section and a single-slotted flap system. Also, boundary-layer transition effects were examined, a segmented leading-edge droop for improved stall/spin resistance was studied, and two roll-controlled devices were evaluated. The wind-tunnel investigation showed that deployment of single-slotted, trailing-edge flap was effective in providing substantial increments in lift required for takeoff and landing performance. Fixed-transition studies to investigate premature tripping of the boundary layer indicated no adverse effects in lift and pitching-moment characteristics for either the cruise or landing configuration. The full-scale results also suggested the need to further optimize the leading-edge droop design that was developed in the subscale tests.

Evaluation of Load Analysis Methods for NASAs GIII Adaptive Compliant Trailing Edge Project

NASA Technical Reports Server (NTRS)

Cruz, Josue; Miller, Eric J.

2016-01-01

The Air Force Research Laboratory (AFRL), NASA Armstrong Flight Research Center (AFRC), and FlexSys Inc. (Ann Arbor, Michigan) have collaborated to flight test the Adaptive Compliant Trailing Edge (ACTE) flaps. These flaps were installed on a Gulfstream Aerospace Corporation (GAC) GIII aircraft and tested at AFRC at various deflection angles over a range of flight conditions. External aerodynamic and inertial load analyses were conducted with the intention to ensure that the change in wing loads due to the deployed ACTE flap did not overload the existing baseline GIII wing box structure. The objective of this paper was to substantiate the analysis tools used for predicting wing loads at AFRC. Computational fluid dynamics (CFD) models and distributed mass inertial models were developed for predicting the loads on the wing. The analysis tools included TRANAIR (full potential) and CMARC (panel) models. Aerodynamic pressure data from the analysis codes were validated against static pressure port data collected in-flight. Combined results from the CFD predictions and the inertial load analysis were used to predict the normal force, bending moment, and torque loads on the wing. Wing loads obtained from calibrated strain gages installed on the wing were used for substantiation of the load prediction tools. The load predictions exhibited good agreement compared to the flight load results obtained from calibrated strain gage measurements.

Tests of a Full-Scale Model of the Republic XF-91 Airplane in the Ames 40- by 80-Foot Wind Tunnel. Force and Moment Data

NASA Technical Reports Server (NTRS)

Hunten, Lynn W.; Dew, Joseph K.

1949-01-01

Wind-tunnel tests of a full-scale model of the Republic XF-91 airplane having swept-back wings and a vee tail were conducted to determine both the stability and control characteristics of the model longitudinally, laterally, and directionally. Configurations of the model were investigated involving such variables as external fuel tanks, a landing gear, trailing-edge flaps, leading-edge slats, and a range of wing incidences and tail incidences.

LES tests on airfoil trailing edge serration

NASA Astrophysics Data System (ADS)

Zhu, Wei Jun; Shen, Wen Zhong

2016-09-01

In the present study, a large number of acoustic simulations are carried out for a low noise airfoil with different Trailing Edge Serrations (TES). The Ffowcs Williams-Hawkings (FWH) acoustic analogy is used for noise prediction at trailing edge. The acoustic solver is running on the platform of our in-house incompressible flow solver EllipSys3D. The flow solution is first obtained from the Large Eddy Simulation (LES), the acoustic part is then carried out based on the instantaneous hydrodynamic pressure and velocity field. To obtain the time history data of sound pressure, the flow quantities are integrated around the airfoil surface through the FWH approach. For all the simulations, the chord based Reynolds number is around 1.5x106. In the test matrix, the effects from angle of attack, the TE flap angle, the length/width of the TES are investigated. Even though the airfoil under investigation is already optimized for low noise emission, most numerical simulations and wind tunnel experiments show that the noise level is further decreased by adding the TES device.

Upper surface blowing noise of the NASA-Ames quiet short-haul research aircraft

NASA Technical Reports Server (NTRS)

Bohn, A. J.; Shovlin, M. D.

1980-01-01

An experimental study of the propulsive-lift noise of the NASA-Ames quiet short-haul research aircraft (QSRA) is described. Comparisons are made of measured QSRA flyover noise and model propulsive-lift noise data available in references. Developmental tests of trailing-edge treatments were conducted using sawtooth-shaped and porous USB flap trailing-edge extensions. Small scale parametric tests were conducted to determine noise reduction/design relationships. Full-scale static tests were conducted with the QSRA preparatory to the selection of edge treatment designs for flight testing. QSRA flight and published model propulsive-lift noise data have similar characteristics. Noise reductions of 2 to 3 dB were achieved over a wide range of frequency and directivity angles in static tests of the QSRA. These noise reductions are expected to be achieved or surpassed in flight tests planned by NASA in 1980.

Small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts. Volume 1. Aerodynamic test results

NASA Technical Reports Server (NTRS)

Renselaer, D. J.; Nishida, R. S.; Wilkin, C. A.

1975-01-01

The results and analyses of aerodynamic and acoustic studies conducted on the small scale noise and wind tunnel tests of upper surface blowing nozzle flap concepts are presented. Various types of nozzle flap concepts were tested. These are an upper surface blowing concept with a multiple slot arrangement with seven slots (seven slotted nozzle), an upper surface blowing type with a large nozzle exit at approximately mid-chord location in conjunction with a powered trailing edge flap with multiple slots (split flow or partially slotted nozzle). In addition, aerodynamic tests were continued on a similar multi-slotted nozzle flap, but with 14 slots. All three types of nozzle flap concepts tested appear to be about equal in overall aerodynamic performance but with the split flow nozzle somewhat better than the other two nozzle flaps in the landing approach mode. All nozzle flaps can be deflected to a large angle to increase drag without significant loss in lift. The nozzle flap concepts appear to be viable aerodynamic drag modulation devices for landing.

Downwash and Wake Behind Plain and Flapped Airfoils

NASA Technical Reports Server (NTRS)

Silverstein, Abe; Katzoff, S; Bullivant, W Kenneth

1939-01-01

Extensive experimental measurements have been made of the downwash angles and the wake characteristics behind airfoils with and without flaps and the data have been analyzed and correlated with the theory. A detailed study was made of the errors involved in applying lifting-line theory, such as the effects of a finite wing chord, the rolling-up of the trailing vortex sheet, and the wake. The downwash angles, as computed from the theoretical span load distribution by means of the Biot-Savart equation, were found to be in satisfactory agreement with the experimental results. The rolling-up of the trailing vortex sheet may be neglected, but the vertical displacement of the vortex sheet requires consideration. By the use of a theoretical treatment indicated by Prandtl, it has been possible to generalize the available experimental results so the predictions can be made of the important wake parameters in terms of the distance behind the airfoil trailing edge and the profile-drag coefficient. The method of application of the theory to design and the satisfactory agreement between predicted and experimental results when applied to an airplane are demonstrated.

Static and Dynamic Aeroelastic Tailoring With Variable Camber Control

NASA Technical Reports Server (NTRS)

Stanford, Bret K.

2016-01-01

This paper examines the use of a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for aeroservoelastic optimization of a transport wingbox. The quasisteady and unsteady motions of the flap system are utilized as design variables, along with patch-level structural variables, towards minimizing wingbox weight via maneuver load alleviation and active flutter suppression. The resulting system is, in general, very successful at removing structural weight in a feasible manner. Limitations to this success are imposed by including load cases where the VCCTEF system is not active (open-loop) in the optimization process, and also by including actuator operating cost constraints.

A Non-linear Lifting Line Model for Design and Analysis of Trochoidal Propulsors

NASA Astrophysics Data System (ADS)

Roesler, Bernard; Epps, Brenden

2014-11-01

Flapping wing propulsors may increase the propulsive efficiency of large shipping vessels. A comparison of the design of a notional propulsor for a large shipping vessel with (a) a conventional ducted propeller versus (b) a flapping wing propulsor is presented. Calculations for flapping wing propulsors are performed using an open-source MATLAB software suite developed by the authors, CyROD, implementing an unsteady lifting-line model with free vortex wake roll-up to study the non-linear effects of foil-wake, and foil-foil interactions. Improvements to the traditional lifting line theory are made using further discretization of the wake vortex ring spacing near the trailing edge. Considerations of packaging options for a flapping wing propulsor on a large shipping vessel are presented, and compared with those for a conventional ducted propeller.

Low-speed wind-tunnel investigation of a large scale advanced arrow-wing supersonic transport configuration with engines mounted above wing for upper-surface blowing

NASA Technical Reports Server (NTRS)

Shivers, J. P.; Mclemore, H. C.; Coe, P. L., Jr.

1976-01-01

Tests have been conducted in a full scale tunnel to determine the low speed aerodynamic characteristics of a large scale advanced arrow wing supersonic transport configuration with engines mounted above the wing for upper surface blowing. Tests were made over an angle of attack range of -10 deg to 32 deg, sideslip angles of + or - 5 deg, and a Reynolds number range of 3,530,000 to 7,330,000. Configuration variables included trailing edge flap deflection, engine jet nozzle angle, engine thrust coefficient, engine out operation, and asymmetrical trailing edge boundary layer control for providing roll trim. Downwash measurements at the tail were obtained for different thrust coefficients, tail heights, and at two fuselage stations.

Transonic Free-Flight Investigation of the Longitudinal Aerodynamic Characteristics of a 1/10-Scale Steel-Wing Model of the Northrop MX-775A Missile with Leading-Edge Extensions, Inboard Trailing-Edge Flaps, and a Speed Brake on the Vertical Tail

NASA Technical Reports Server (NTRS)

Arbic, R. G.

1955-01-01

Results are presented of a free-flight investigation between Mach numbers of 0.7 to 1.3 and Reynolds numbers of 3.1 x 10(exp 6) to 7.0 x 10(exp 6) to determine the longitudinal aerodynamic characteristics of the Northrop MX-775A missile. This missile has a weng, body, and vertical tail, but has no horizontal tail. The basic wing plan form has an aspect ratio of 5.5, 45 deg of sweepback of the 0.406 streamwise chord line, and a taper ratio of 0.4. A 1/10-scale steel-wing model of the missile was flown with modifications to the basic wing plan form consisting of leading-edge chord-extensions deflected 7 deg downward together with the forward 15 percent of the wing chord, and inboard trailing-edge flaps deflected 5 deg downward. In addition, the model had a static-pressure tube mounted at the tip of the vertical tail for position-error measurements and had a speed brake also mounted on the vertical tail to trim the model to positive lift coefficients and to permit determination of the trim and drag effectiveness of the brake. The data are uncorrected for the effects of wing elasticity, but experimental wing influence coefficients are presented.

Two dimensional aerodynamic interference effects on oscillating airfoils with flaps in ventilated subsonic wind tunnels. [computational fluid dynamics

NASA Technical Reports Server (NTRS)

Fromme, J.; Golberg, M.; Werth, J.

1979-01-01

The numerical computation of unsteady airloads acting upon thin airfoils with multiple leading and trailing-edge controls in two-dimensional ventilated subsonic wind tunnels is studied. The foundation of the computational method is strengthened with a new and more powerful mathematical existence and convergence theory for solving Cauchy singular integral equations of the first kind, and the method of convergence acceleration by extrapolation to the limit is introduced to analyze airfoils with flaps. New results are presented for steady and unsteady flow, including the effect of acoustic resonance between ventilated wind-tunnel walls and airfoils with oscillating flaps. The computer program TWODI is available for general use and a complete set of instructions is provided.

76 FR 61633 - Airworthiness Directives; Bombardier, Inc. Model CL-600-2C10 (Regional Jet Series 700, 701, & 702...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-10-05

.... (6) The climb ceiling obtained from the Flight Planning and Cruise Control Manual (FPCCM) must be... from the Flight Planning and Cruise Control Manual (FPCCM) must be reduced by 1,000 ft/door.'' Note 4... the landing phase of flight. The door damaged the trailing edge flap and punctured the rear fuselage...

Wind tunnel test of a tailless aircraft with a belly-flap control surface used in combination with wing flaps

NASA Astrophysics Data System (ADS)

Dougherty, Daniel A.

A wind tunnel tat of a tailless aircraft configuration that has been quipped with a belly-flap control surface, was conducted with the goal of improving the trimmed maximum-lift coefficient. Tailless aircraft have aerodynamic and structural efficiencies that are superior to those of a traditionally configured wing/body/tail aircraft. However, tailless aircraft have a low maximum-lift coefficient such that; when sized for equivalent takeoff performance, the tailless aircraft suffers a large reduction in aerodynamic and structural efficiencies. A Belly-Flap control surface used in combination with wing trailing edge flaps was tested in a wind tunnel with the goal of achieving a longitudinally trimmed solution at a higher maximum lift coefficient. It was determined that, though the Belly-Flap increases the trimmed lift of the tailless configuration at low angles of attack, the maximum lift coefficient is slightly reduced in relation to the controls neutral configuration.

Coupled Vortex-Lattice Flight Dynamic Model with Aeroelastic Finite-Element Model of Flexible Wing Transport Aircraft with Variable Camber Continuous Trailing Edge Flap for Drag Reduction

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric; Nguyen, Daniel; Dao, Tung; Trinh, Khanh

2013-01-01

This paper presents a coupled vortex-lattice flight dynamic model with an aeroelastic finite-element model to predict dynamic characteristics of a flexible wing transport aircraft. The aircraft model is based on NASA Generic Transport Model (GTM) with representative mass and stiffness properties to achieve a wing tip deflection about twice that of a conventional transport aircraft (10% versus 5%). This flexible wing transport aircraft is referred to as an Elastically Shaped Aircraft Concept (ESAC) which is equipped with a Variable Camber Continuous Trailing Edge Flap (VCCTEF) system for active wing shaping control for drag reduction. A vortex-lattice aerodynamic model of the ESAC is developed and is coupled with an aeroelastic finite-element model via an automated geometry modeler. This coupled model is used to compute static and dynamic aeroelastic solutions. The deflection information from the finite-element model and the vortex-lattice model is used to compute unsteady contributions to the aerodynamic force and moment coefficients. A coupled aeroelastic-longitudinal flight dynamic model is developed by coupling the finite-element model with the rigid-body flight dynamic model of the GTM.

Estimated Benefits of Variable-Geometry Wing Camber Control for Transport Aircraft

NASA Technical Reports Server (NTRS)

Bolonkin, Alexander; Gilyard, Glenn B.

1999-01-01

Analytical benefits of variable-camber capability on subsonic transport aircraft are explored. Using aerodynamic performance models, including drag as a function of deflection angle for control surfaces of interest, optimal performance benefits of variable camber are calculated. Results demonstrate that if all wing trailing-edge surfaces are available for optimization, drag can be significantly reduced at most points within the flight envelope. The optimization approach developed and illustrated for flight uses variable camber for optimization of aerodynamic efficiency (maximizing the lift-to-drag ratio). Most transport aircraft have significant latent capability in this area. Wing camber control that can affect performance optimization for transport aircraft includes symmetric use of ailerons and flaps. In this paper, drag characteristics for aileron and flap deflections are computed based on analytical and wind-tunnel data. All calculations based on predictions for the subject aircraft and the optimal surface deflection are obtained by simple interpolation for given conditions. An algorithm is also presented for computation of optimal surface deflection for given conditions. Benefits of variable camber for a transport configuration using a simple trailing-edge control surface system can approach more than 10 percent, especially for nonstandard flight conditions. In the cruise regime, the benefit is 1-3 percent.

Smart wing rotation and trailing-edge vortices enable high frequency mosquito flight

PubMed Central

Bomphrey, Richard J.; Nakata, Toshiyuki; Phillips, Nathan; Walker, Simon M.

2017-01-01

Summary Mosquitoes exhibit unique wing kinematics; their long, slender wings flap at remarkably high frequencies for their size (>800 Hz) and with lower stroke amplitudes than any other insect group1. This shifts weight support away from the translation-dominated, aerodynamic mechanisms used by most insects2, as well as by helicopters and aeroplanes, towards poorly understood rotational mechanisms that occur when pitching at the end of each half-stroke. Here we report wing kinematics and solve the full Navier-Stokes equations using computational fluid dynamics with overset grids and validate our results with in vivo flow measurements. We show that, while familiar separated flow patterns are used by mosquitoes, much of the aerodynamic force that supports their weight is generated in a manner unlike any previously described flying animal. In total, there are three key features: leading-edge vortices (a well-known mechanism that appears to be almost ubiquitous in insect flight), trailing-edge vortices caused by a novel form of wake capture at stroke reversal, and rotational drag. The two new elements are largely independent of the wing velocity, instead relying on rapid changes in the pitch angle (wing rotation) at the end of each half stroke, and are therefore relatively immune to the shallow flapping amplitude. Moreover, these mechanisms are particularly well-suited to high-aspect ratio mosquito wings. PMID:28355184

Smart wing rotation and trailing-edge vortices enable high frequency mosquito flight

NASA Astrophysics Data System (ADS)

Bomphrey, Richard J.; Nakata, Toshiyuki; Phillips, Nathan; Walker, Simon M.

2017-03-01

Mosquitoes exhibit unusual wing kinematics; their long, slender wings flap at remarkably high frequencies for their size (>800 Hz)and with lower stroke amplitudes than any other insect group. This shifts weight support away from the translation-dominated, aerodynamic mechanisms used by most insects, as well as by helicopters and aeroplanes, towards poorly understood rotational mechanisms that occur when pitching at the end of each half-stroke. Here we report free-flight mosquito wing kinematics, solve the full Navier-Stokes equations using computational fluid dynamics with overset grids, and validate our results with in vivo flow measurements. We show that, although mosquitoes use familiar separated flow patterns, much of the aerodynamic force that supports their weight is generated in a manner unlike any previously described for a flying animal. There are three key features: leading-edge vortices (a well-known mechanism that appears to be almost ubiquitous in insect flight), trailing-edge vortices caused by a form of wake capture at stroke reversal, and rotational drag. The two new elements are largely independent of the wing velocity, instead relying on rapid changes in the pitch angle (wing rotation) at the end of each half-stroke, and they are therefore relatively immune to the shallow flapping amplitude. Moreover, these mechanisms are particularly well suited to high aspect ratio mosquito wings.

Wind-Tunnel Investigation of the Horizontal Motion of a Wing Near the Ground

NASA Technical Reports Server (NTRS)

Serebrisky, Y. M.; Biachuev, S. A.

1946-01-01

By the method of images the horizontal steady motion of a wing at small heights above the ground was investigated in the wind tunnel, A rectangular wing with Clark Y-H profile was tested with and without flaps. The distance from the trailing edge of the wing to the ground was varied within the limits 0.75 less than or = s/c less than or = 0.25. Measurements were made of the lift, the drag, the pitching moment, and the pressure distribution at one section. For a wing without flaps and one with flaps a considereble decrease in the lift force and a,drop in the drag was obtained at angles of attack below stalling. The flow separation near the ground occurs at smaller angles of attack than is the case for a great height above the ground. At horizontal steady flight for practical values of the height above the ground the maximum lift coefficient for the wing without flaps changes little, but markedly decreases for the wing with flaps. Analysis of these phenomena involves the investigation of the pressure distribution. The pressure distribution curves showed that the changes occurring near the ground are not equivalent to a change in the angle of attack. At the lower surface of the section a very strong increase in the pressures is observed. The pressure changes on the upper surface at angles of attack below stalling are insignificant and lead mainly to an increase in the unfavorable pressure gradient, resulting in the earlier occurrence of separation. For a wing with flaps at large angles of attack for distances from the trailing edge of the flap to the ground less than 0.5 chord, the flow between the wing end the ground is retarded so greatly that the pressure coefficient at the lower surface of the section is very near its limiting value (P = 1), and any further possibility of increase in the pressure is very small. In the application an approximate computation procedure is given of the change of certain aerodynamic characteristics for horizontal steady flight near the ground.

Longitudinal Stability and Control Characteristics as Determined by the Rocket-model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-aspect-ratio Wing Having Trailing-edge Flap Controls for a Mach Number Range of 0.7 to 1.

NASA Technical Reports Server (NTRS)

Baber, Hal T , Jr; Moul, Martin T

1955-01-01

Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle - of-attack range of this test (0 deg to 8 deg). The aerodynamic-center location for angles of attack near 50 remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near 0 deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of 0 deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle -of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Longitudinal Stability and Control Characteristics as Determined by the Rocket-Model Technique for an Inline, Cruciform, Canard Missile Configuration with a Low-Aspect-Ratio Wing Having Trailing-Edge Flap Controls for a Mach Number Range of 0.7 to 1.8

NASA Technical Reports Server (NTRS)

Baber, H. T., Jr.; Moul, M. T.

1955-01-01

Two full-scale models of an inline, cruciform, canard missile configuration having a low-aspect-ratio wing equipped with flap-type controls were flight tested in order to determine the missile's longitudinal aerodynamic characteristics. Stability derivatives and control and drag characteristics are presented for a range of Mach number from 0.7 to 1.8. Nonlinear lift and moment curves were noted for the angle-of-attack range of this test (0 deg to 8 deg ). The aerodynamic-center location for angles of attack near 5 deg remained nearly constant for supersonic speeds at 13.5 percent of the mean aerodynamic chord; whereas for angles of attack near O deg, there was a rapid forward movement of the aerodynamic center as the Mach number increased. At a control deflection of O deg, the missile's response to the longitudinal control was in an essentially fixed space plane which was not coincident with the pitch plane as a result of the missile rolling. As a consequence, stability characteristics were determined from the resultant of pitch and yaw motions. The damping-in-pitch derivatives for the two angle-of-attack ranges of the test are in close agreement and varied only slightly with Mach number. The horn-balanced trailing-edge flap was effective in producing angle of attack over the Mach number range.

Stability and performance characteristics of a fixed arrow wing supersonic transport configuration (SCAT 15F-9898) at Mach numbers from 0.60 to 1.20

NASA Technical Reports Server (NTRS)

Decker, J. P.; Jacobs, P. F.

1978-01-01

Tests on a 0.015 scale model of a supersonic transport were conducted at Mach numbers from 0.60 to 1.20. Tests of the complete model with three wing planforms, two different leading-edge radii, and various combinations of component parts, including both leading- and trailing-edge flaps, were made over an angle-of-attack range from about -6 deg to 13 deg and at sideslip angles of 0 deg and 2 deg.

An analytical model and scaling of chordwise flexible flapping wings in forward flight.

PubMed

Kodali, Deepa; Kang, Chang-Kwon

2016-12-13

Aerodynamic performance of biological flight characterized by the fluid structure interaction of a flapping wing and the surrounding fluid is affected by the wing flexibility. One of the main challenges to predict aerodynamic forces is that the wing shape and motion are a priori unknown. In this study, we derive an analytical fluid-structure interaction model for a chordwise flexible flapping two-dimensional airfoil in forward flight. A plunge motion is imposed on the rigid leading-edge (LE) of teardrop shape and the flexible tail dynamically deforms. The resulting unsteady aeroelasticity is modeled with the Euler-Bernoulli-Theodorsen equation under a small deformation assumption. The two-way coupling is realized by considering the trailing-edge deformation relative to the LE as passive pitch, affecting the unsteady aerodynamics. The resulting wing deformation and the aerodynamic performance including lift and thrust agree well with high-fidelity numerical results. Under the dynamic balance, the aeroelastic stiffness decreases, whereas the aeroelastic stiffness increases with the reduced frequency. A novel aeroelastic frequency ratio is derived, which scales with the wing deformation, lift, and thrust. Finally, the dynamic similarity between flapping in water and air is established.

Power performance optimization and loads alleviation with active flaps using individual flap control

NASA Astrophysics Data System (ADS)

Pettas, Vasilis; Barlas, Thanasis; Gertz, Drew; Madsen, Helge A.

2016-09-01

The present article investigates the potential of Active Trailing Edge Flaps (ATEF) in terms of increase in annual energy production (AEP) as well as reduction of fatigue loads. The basis for this study is the DTU 10 MW Reference Wind Turbine (RWT) simulated using the aeroelastic code HAWC2. In an industrial-oriented manner the baseline rotor is upscaled by 5% and the ATEFs are implemented in the outer 30% of the blades. The flap system is kept simple and robust with a single flap section and control with wind speed, rotor azimuth, root bending moments and angle of attack in flap's mid-section being the sensor inputs. The AEP is increased due to the upscaling but also further due to the flap system while the fatigue loads in components of interest (blade, tower, nacelle and main bearing) are reduced close to the level of the original turbine. The aim of this study is to demonstrate a simple and applicable method that can be a technology enabler for rotor upscaling and lowering cost of energy.

Transonic Free-To-Roll Analysis of the F/A-18E and F-35 Configurations

NASA Technical Reports Server (NTRS)

Owens, D. Bruce; McConnell, Jeffrey K.; Brandon, Jay M.; Hall, Robert M.

2004-01-01

The free-to-roll technique is used as a tool for predicting areas of uncommanded lateral motions. Recently, the NASA/Navy/Air Force Abrupt Wing Stall Program extended the use of this technique to the transonic speed regime. Using this technique, this paper evaluates various wing configurations on the pre-production F/A-18E aircraft and the Joint Strike Fighter (F-35) aircraft. The configurations investigated include leading and trailing edge flap deflections, fences, leading edge flap gap seals, and vortex generators. These tests were conducted in the NASA Langley 16-Foot Transonic Tunnel. The analysis used a modification of a figure-of-merit developed during the Abrupt Wing Stall Program to discern configuration effects. The results showed how the figure-of-merit can be used to schedule wing flap deflections to avoid areas of uncommanded lateral motion. The analysis also used both static and dynamic wind tunnel data to provide insight into the uncommanded lateral behavior. The dynamic data was extracted from the time history data using parameter identification techniques. In general, modifications to the pre-production F/A-18E resulted in shifts in angle-of-attack where uncommanded lateral activity occurred. Sealing the gap between the inboard and outboard leading-edge flaps on the Navy version of the F-35 eliminated uncommanded lateral activity or delayed the activity to a higher angle-of-attack.

NACA Conference on Some Problems of Aircraft Operation Held at Lewis Flight Propulsion Lab., Cleveland, Ohio on October 9-10, 1950

DTIC Science & Technology

1950-10-10

1947. 91 20. Salmi, Reino , J.: Effects of Leading-Edge Devices and Trailing- F,4,e Flaps on Longitudinal Characteristics of Two 47.70 Swept- back...in a similar manner, but the entire handling qualities specifications -re too numerous to allow dio - cussion in this paper. In order to illust-,ate the

Low-speed wind tunnel investigation of a semispan STOL jet transport wing body with an upper surface blown jet flap

NASA Technical Reports Server (NTRS)

Phelps, A. E., III; Letko, W.; Henderson, R. L.

1973-01-01

An investigation of the static longitudinal aerodynamic characteristics of a semispan STOL jet transport wing-body with an upper-surface blown jet flap for lift augmentation was conducted in a low-speed wind tunnel having a 12-ft octagonal test section. The semispan swept wing had an aspect ratio of 3.92 (7.84 for the full span) and had two simulated turbofan engines mounted ahead of and above the wing in a siamese pod equipped with an exhaust deflector. The purpose of the deflector was to spread the engine exhaust into a jet sheet attached to the upper surface of the wing so that it would turn downward over the flap and provide lift augmentation. The wing also had optional boundary-layer control provided by air blowing through a thin slot over a full-span plain trailing-edge flap.

Pressure distributions on a rectangular aspect-ratio-6, slotted supercritical airfoil wing with externally blown flaps

NASA Technical Reports Server (NTRS)

Johnson, W. G., Jr.

1976-01-01

An investigation was made in the 5.18 m (17 ft) test section of the Langley 300 MPH 7 by 10 foot tunnel on a rectangular, aspect ratio 6 wing which had a slotted supercritical airfoil section and externally blown flaps. The 13 percent thick wing was fitted with two high lift flap systems: single slotted and double slotted. The designations single slotted and double slotted do not include the slot which exists near the trailing edge of the basic slotted supercritical airfoil. Tests were made over an angle of attack range of -6 deg to 20 deg and a thrust-coefficient range up to 1.94 for a free-stream dynamic pressure of 526.7 Pa (11.0 lb/sq ft). The results of the investigation are presented as curves and tabulations of the chordwise pressure distributions at the midsemispan station for the wing and each flap element.

Augmentation of maneuver performance by spanwise blowing

NASA Technical Reports Server (NTRS)

Erickson, G. E.; Campbell, J. F.

1977-01-01

A generalized wind tunnel model was tested to investigate new component concepts utilizing spanwise blowing to provide improved maneuver characteristics for advanced fighter aircraft. Primary emphasis was placed on high angle of attack performance, stability, and control at subsonic speeds. Spanwise blowing on a 44 deg swept trapezoidal wing resulted in leading edge vortex enhancement with subsequent large vortex-induced lift increments and drag polar improvements at the higher angles of attack. Small deflections of a leading edge flap delayed these lift and drag benefits to higher angles of attack. In addition, blowing was more effective at higher Mach numbers. Spanwise blowing in conjunction with a deflected trailing edge flap resulted in lift and drag benefits that exceeded the summation of the effects of each high lift device acting alone. Asymmetric blowing was an effective lateral control device at the higher angles of attack. Spanwise blowing on the wing reduced horizontal tail loading and improved the lateral-directional stability characteristics of a wing-horizontal tail-vertical tail configuration.

Blade vortex interaction noise reduction techniques for a rotorcraft

NASA Technical Reports Server (NTRS)

Charles, Bruce D. (Inventor); Hassan, Ahmed A. (Inventor); Tadghighi, Hormoz (Inventor); JanakiRam, Ram D. (Inventor); Sankar, Lakshmi N. (Inventor)

1996-01-01

An active control device for reducing blade-vortex interactions (BVI) noise generated by a rotorcraft, such as a helicopter, comprises a trailing edge flap located near the tip of each of the rotorcraft's rotor blades. The flap may be actuated in any conventional way, and is scheduled to be actuated to a deflected position during rotation of the rotor blade through predetermined regions of the rotor azimuth, and is further scheduled to be actuated to a retracted position through the remaining regions of the rotor azimuth. Through the careful azimuth-dependent deployment and retraction of the flap over the rotor disk, blade tip vortices which are the primary source for BVI noise are (a) made weaker and (b) pushed farther away from the rotor disk (that is, larger blade-vortex separation distances are achieved).

Upper-surface-blowing flow-turning performance

NASA Technical Reports Server (NTRS)

Sleeman, W. C., Jr.; Phelps, A. E., III

1976-01-01

Jet exhaust flow-turning characteristics were determined for systematic variations in upper-surface blowing exhaust nozzles and trailing-edge flap configuration variables from experimental wind-off (static) flow studies. For conditions with parallel flow exhausting from the nozzle, jet height (as indicated by nozzle exit height) and flap radius were found to be the most important parameters relating to flow turning. Nonparallel flow from the nozzle, as obtained from an internal roof angle and/or side spread angle, had a large favorable effect on flow turning. Comparisons made between static turning results and wind tunnel aerodynamic studies of identical configurations indicated that static flow-turning results can be indicative of wind-on powered lift performance for both good and poor nozzle-flap combinations but, for marginal designs, can lead to overly optimistic assessment of powered lift potential.

Blade vortex interaction noise reduction techniques for a rotorcraft

NASA Technical Reports Server (NTRS)

Charles, Bruce D. (Inventor); JanakiRam, Ram D. (Inventor); Hassan, Ahmed A. (Inventor); Tadghighi, Hormoz (Inventor); Sankar, Lakshmi N. (Inventor)

1998-01-01

An active control device for reducing blade-vortex interactions (BVI) noise generated by a rotorcraft, such as a helicopter, comprises a trailing edge flap located near the tip of each of the rotorcraft's rotor blades. The flap may be actuated in any conventional way, and is scheduled to be actuated to a deflected position during rotation of the rotor blade through predetermined regions of the rotor azimuth, and is further scheduled to be actuated to a retracted position through the remaining regions of the rotor azimuth. Through the careful azimuth-dependent deployment and retraction of the flap over the rotor disk, blade tip vortices which are the primary source for BVI noise are (a) made weaker and (b) pushed farther away from the rotor disk (that is, larger blade-vortex separation distances are achieved).

Aeroelastic Wing Shaping Control Subject to Actuation Constraints.

NASA Technical Reports Server (NTRS)

Swei, Sean Shan-Min; Nguyen, Nhan

2014-01-01

This paper considers the control of coupled aeroelastic aircraft model which is configured with Variable Camber Continuous Trailing Edge Flap (VCCTEF) system. The relative deflection between two adjacent flaps is constrained and this actuation constraint is accounted for when designing an effective control law for suppressing the wing vibration. A simple tuned-mass damper mechanism with two attached masses is used as an example to demonstrate the effectiveness of vibration suppression with confined motion of tuned masses. In this paper, a dynamic inversion based pseudo-control hedging (PCH) and bounded control approach is investigated, and for illustration, it is applied to the NASA Generic Transport Model (GTM) configured with VCCTEF system.

Computational studies of the effects of active and passive circulation enhancement concepts on wind turbine performance

NASA Astrophysics Data System (ADS)

Tongchitpakdee, Chanin

With the advantage of modern high speed computers, there has been an increased interest in the use of first-principles based computational approaches for the aerodynamic modeling of horizontal axis wind turbine (HAWT). Since these approaches are based on the laws of conservation (mass, momentum, and energy), they can capture much of the physics in great detail. The ability to accurately predict the airloads and power output can greatly aid the designers in tailoring the aerodynamic and aeroelastic features of the configuration. First-principles based analyses are also valuable for developing active means (e.g., circulation control), and passive means (e.g., Gurney flaps) of reducing unsteady blade loads, mitigating stall, and for efficient capture of wind energy leading to more electrical power generation. In this present study, the aerodynamic performance of a wind turbine rotor equipped with circulation enhancement technology (trailing edge blowing or Gurney flaps) is investigated using a three-dimensional unsteady viscous flow analysis. The National Renewable Energy Laboratory (NREL) Phase VI horizontal axis wind turbine is chosen as the baseline configuration. Prior to its use in exploring these concepts, the flow solver is validated with the experimental data for the baseline case under yawed flow conditions. Results presented include radial distribution of normal and tangential forces, shaft torque, root flap moment, surface pressure distributions at selected radial locations, and power output. Results show that good agreement has been for a range of wind speeds and yaw angles, where the flow is attached. At high wind speeds, however, where the flow is fully separated, it was found that the fundamental assumptions behind this present methodology breaks down for the baseline turbulence model (Spalart-Allmaras model), giving less accurate results. With the implementation of advanced turbulence model, Spalart-Allmaras Detached Eddy Simulation (SA-DES), the accuracy of the results at high wind speeds are improved. Results of circulation enhancement concepts show that, at low wind speed (attached flow) conditions, a Coanda jet at the trailing edge of the rotor blade is effective at increasing circulation resulting in an increase of lift and the chordwise thrust force. This leads to an increased amount of net power generation compared to the baseline configuration for moderate blowing coefficients. The effects of jet slot height and pulsed jet are also investigated in this study. A passive Gurney flap was found to increase the bound circulation and produce increased power in a manner similar to the Coanda jet. At high wind speed where the flow is separated, both the Coanda jet and Gurney flap become ineffective. Results of leading edge blowing indicate that a leading edge blowing jet is found to be beneficial in increasing power generation at high wind speeds. The effect of Gurney flap angle is also studied. Gurney flap angle has significant influence in power generation. Higher power output is obtained at higher flap angles.

Full Scale Span Load Distribution on a Tapered Wing with Split Flaps of Various Spans

NASA Technical Reports Server (NTRS)

Parsons, John F; Silverstein, Abe

1937-01-01

Pressure-distribution tests were conducted in the full-scale wind tunnel on a 2:1 tapered U.S.A. 45 airfoil equipped with 20 percent chord split trailing-edge flaps of various spans. A special installation was employed in the tests utilizing a half-span airfoil mounted vertically above a reflection plane. The airfoil has a constant chord center section and rounded tips and is tapered in thickness from 18 percent c at the root to 9 percent c at the tip. The aerodynamic characteristics, given by the usual dimension less coefficients, are presented graphically as functions of flap span and angle of attack as well as by semispan load diagrams. The results indicate, in general, that only a relatively small increase in the normal-force coefficient is to be expected by extending the flap span of an airfoil-flap combination, similar to the one tested, beyond 70 percent of the wing span.

Static, noise, and transition tests of a combined-surface-blowing V/STOL lift/propulsion system

NASA Technical Reports Server (NTRS)

Schoen, A. H.; Kolesar, C. E.; Schaeffer, E. G.

1977-01-01

Efficient thrust vectoring and high levels of circulatory lift were obtained in tests of a half model V/STOL airplane by using a type of externally blown jet flap in which the jet exhaust from wing-mounted cruise fans is directed over both upper and lower surfaces of a flapped wing. Approximately 90% thrust recovery with 87 deg of thrust vectoring was achieved under static conditions using 89 deg of trailing edge flap deflection. The approximately 10% loss appears to be associated primarily with pressure losses due to the flap brackets or slot entries. The jet induced lift was shown to be 55% of the theoretical value for a fullspan jet-flapped wing, even though only 27.5% of the wingspan was immersed in the jet. Steady rate of descent capability in excess of 1,000 feet per minute is predicted. The possibility of significant aerodynamic-noise cancelling when blowing over both surfaces at high velocities is indicated.

The investigation of a variable camber blade lift control for helicopter rotor systems

NASA Technical Reports Server (NTRS)

Awani, A. O.

1982-01-01

A new rotor configuration called the variable camber rotor was investigated numerically for its potential to reduce helicopter control loads and improve hover performance. This rotor differs from a conventional rotor in that it incorporates a deflectable 50% chord trailing edge flap to control rotor lift, and a non-feathering (fixed) forward portion. Lift control is achieved by linking the blade flap to a conventional swashplate mechanism; therefore, it is pilot action to the flap deflection that controls rotor lift and tip path plane tilt. This report presents the aerodynamic characteristics of the flapped and unflapped airfoils, evaluations of aerodynamics techniques to minimize flap hinge moment, comparative hover rotor performance and the physical concepts of the blade motion and rotor control. All the results presented herein are based on numerical analyses. The assessment of payoff for the total configuration in comparison with a conventional blade, having the same physical characteristics as an H-34 helicopter rotor blade was examined for hover only.

Morphing Wings: A Study Using High-Fidelity Aerodynamic Shape Optimization

NASA Astrophysics Data System (ADS)

Curiale, Nathanael J.

With the aviation industry under pressure to reduce fuel consumption, morphing wings have the capacity to improve aircraft performance, thereby making a significant contribution to reversing climate change. Through high-fidelity aerodynamic shape optimization, various forms of morphing wings are assessed for a hypothetical regional-class aircraft. The framework used solves the Reynolds-averaged Navier-Stokes equations and utilizes a gradient-based optimization algorithm. Baseline geometries are developed through multipoint optimization, where the average drag coefficient is minimized over a range of flight conditions with additional dive constraints. Morphing optimizations are then performed, beginning with these baseline shapes. Five distinct types of morphing are investigated and compared. Overall, a theoretical fully adaptable wing produces roughly a 2% improvement in average performance, whereas trailing-edge morphing with a 27-point multipoint baseline results in just over a 1% improvement in average performance. Trailing-edge morphing proves to be more beneficial than leading-edge morphing, upper-surface morphing, and a conventional flap.

AFC-Enabled Simplified High-Lift System Integration Study

NASA Technical Reports Server (NTRS)

Hartwich, Peter M.; Dickey, Eric D.; Sclafani, Anthony J.; Camacho, Peter; Gonzales, Antonio B.; Lawson, Edward L.; Mairs, Ron Y.; Shmilovich, Arvin

2014-01-01

The primary objective of this trade study report is to explore the potential of using Active Flow Control (AFC) for achieving lighter and mechanically simpler high-lift systems for transonic commercial transport aircraft. This assessment was conducted in four steps. First, based on the Common Research Model (CRM) outer mold line (OML) definition, two high-lift concepts were developed. One concept, representative of current production-type commercial transonic transports, features leading edge slats and slotted trailing edge flaps with Fowler motion. The other CRM-based design relies on drooped leading edges and simply hinged trailing edge flaps for high-lift generation. The relative high-lift performance of these two high-lift CRM variants is established using Computational Fluid Dynamics (CFD) solutions to the Reynolds-Averaged Navier-Stokes (RANS) equations for steady flow. These CFD assessments identify the high-lift performance that needs to be recovered through AFC to have the CRM variant with the lighter and mechanically simpler high-lift system match the performance of the conventional high-lift system. Conceptual design integration studies for the AFC-enhanced high-lift systems were conducted with a NASA Environmentally Responsible Aircraft (ERA) reference configuration, the so-called ERA-0003 concept. These design trades identify AFC performance targets that need to be met to produce economically feasible ERA-0003-like concepts with lighter and mechanically simpler high-lift designs that match the performance of conventional high-lift systems. Finally, technical challenges are identified associated with the application of AFC-enabled highlift systems to modern transonic commercial transports for future technology maturation efforts.

Development of Advanced High Lift Leading Edge Technology for Laminar Flow Wings

NASA Technical Reports Server (NTRS)

Bright, Michelle M.; Korntheuer, Andrea; Komadina, Steve; Lin, John C.

2013-01-01

This paper describes the Advanced High Lift Leading Edge (AHLLE) task performed by Northrop Grumman Systems Corporation, Aerospace Systems (NGAS) for the NASA Subsonic Fixed Wing project in an effort to develop enabling high-lift technology for laminar flow wings. Based on a known laminar cruise airfoil that incorporated an NGAS-developed integrated slot design, this effort involved using Computational Fluid Dynamics (CFD) analysis and quality function deployment (QFD) analysis on several leading edge concepts, and subsequently down-selected to two blown leading-edge concepts for testing. A 7-foot-span AHLLE airfoil model was designed and fabricated at NGAS and then tested at the NGAS 7 x 10 Low Speed Wind Tunnel in Hawthorne, CA. The model configurations tested included: baseline, deflected trailing edge, blown deflected trailing edge, blown leading edge, morphed leading edge, and blown/morphed leading edge. A successful demonstration of high lift leading edge technology was achieved, and the target goals for improved lift were exceeded by 30% with a maximum section lift coefficient (Cl) of 5.2. Maximum incremental section lift coefficients ( Cl) of 3.5 and 3.1 were achieved for a blown drooped (morphed) leading edge concept and a non-drooped leading edge blowing concept, respectively. The most effective AHLLE design yielded an estimated 94% lift improvement over the conventional high lift Krueger flap configurations while providing laminar flow capability on the cruise configuration.

Validation of a computer code for analysis of subsonic aerodynamic performance of wings with flaps in combination with a canard or horizontal tail and an application to optimization

NASA Technical Reports Server (NTRS)

Carlson, Harry W.; Darden, Christine M.; Mann, Michael J.

1990-01-01

Extensive correlations of computer code results with experimental data are employed to illustrate the use of a linearized theory, attached flow method for the estimation and optimization of the longitudinal aerodynamic performance of wing-canard and wing-horizontal tail configurations which may employ simple hinged flap systems. Use of an attached flow method is based on the premise that high levels of aerodynamic efficiency require a flow that is as nearly attached as circumstances permit. The results indicate that linearized theory, attached flow, computer code methods (modified to include estimated attainable leading-edge thrust and an approximate representation of vortex forces) provide a rational basis for the estimation and optimization of aerodynamic performance at subsonic speeds below the drag rise Mach number. Generally, good prediction of aerodynamic performance, as measured by the suction parameter, can be expected for near optimum combinations of canard or horizontal tail incidence and leading- and trailing-edge flap deflections at a given lift coefficient (conditions which tend to produce a predominantly attached flow).

Portable tomographic PIV measurements of swimming shelled Antarctic pteropods

NASA Astrophysics Data System (ADS)

Adhikari, Deepak; Webster, Donald R.; Yen, Jeannette

2016-12-01

A portable tomographic particle image velocimetry (tomographic PIV) system is described. The system was successfully deployed in Antarctica to study shelled Antarctic pteropods ( Limacina helicina antarctica)—a delicate organism with an unusual propulsion mechanism. The experimental setup consists of a free-standing frame assembled with optical rails, thus avoiding the need for heavy and bulky equipment (e.g. an optical table). The cameras, lasers, optics, and tanks are all rigidly supported within the frame assembly. The results indicate that the pteropods flap their parapodia (or "wings") downward during both power and recovery strokes, which is facilitated by the pitching of their shell. Shell pitching significantly alters the flapping trajectory, allowing the pteropod to move vertically and/or horizontally. The pronation and supination of the parapodia, together with the figure-eight motion during flapping, suggest similarities with insect flight. The volumetric velocity field surrounding the freely swimming pteropod reveals the generation of an attached vortex ring connecting the leading-edge vortex to the trailing-edge vortex during power stroke and a presence of a leading-edge vortex during recovery stroke. These vortex structures play a major role in accelerating the organism vertically and indicate that forces generated on the parapodia during flapping constitute both lift and drag. After completing each stroke, two vortex rings are shed into the wake of the pteropod. The complex combination of body kinematics (parapodia flapping, shell pitch, sawtooth trajectory), flow structures, and resulting force balance may be significantly altered by thinning of the pteropod shell, thus making pteropods an indicator of the detrimental effects of ocean acidification.

14 CFR Appendix B to Part 121 - Airplane Flight Recorder Specification

Code of Federal Regulations, 2010 CFR

2010-01-01

... Transmitter Keying On-Off (Discrete) ±2° ±2% Thrust/Power on Each Engine Full Range Forward ±2° 1 (per engine) 0.2% 2 Trailing Edge Flap or Cockpit Control Selection Full Range or Each Discrete Position ±3° or... Discrete Position ±3° or as Pilot's Indicator 0.5 0.5% 2 Thrust Reverser Position Stowed, In Transit, and...

Development of Variable Camber Continuous Trailing Edge Flap for Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Kaul, Upender; Lebofsky, Sonia; Ting, Eric; Chaparro, Daniel; Urnes, James

2015-01-01

This paper summarizes the recent development of an adaptive aeroelastic wing shaping control technology called variable camber continuous trailing edge flap (VCCTEF). As wing flexibility increases, aeroelastic interactions with aerodynamic forces and moments become an increasingly important consideration in aircraft design and aerodynamic performance. Furthermore, aeroelastic interactions with flight dynamics can result in issues with vehicle stability and control. The initial VCCTEF concept was developed in 2010 by NASA under a NASA Innovation Fund study entitled "Elastically Shaped Future Air Vehicle Concept," which showed that highly flexible wing aerodynamic surfaces can be elastically shaped in-flight by active control of wing twist and bending deflection in order to optimize the spanwise lift distribution for drag reduction. A collaboration between NASA and Boeing Research & Technology was subsequently funded by NASA from 2012 to 2014 to further develop the VCCTEF concept. This paper summarizes some of the key research areas conducted by NASA during the collaboration with Boeing Research and Technology. These research areas include VCCTEF design concepts, aerodynamic analysis of VCCTEF camber shapes, aerodynamic optimization of lift distribution for drag minimization, wind tunnel test results for cruise and high-lift configurations, flutter analysis and suppression control of flexible wing aircraft, and multi-objective flight control for adaptive aeroelastic wing shaping control.

Aeroelastic Modeling of Elastically Shaped Aircraft Concept via Wing Shaping Control for Drag Reduction

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; James Urnes, Sr.

2012-01-01

Lightweight aircraft design has received a considerable attention in recent years as a means for improving cruise efficiency. Reducing aircraft weight results in lower lift requirements which directly translate into lower drag, hence reduced engine thrust requirements during cruise. The use of lightweight materials such as advanced composite materials has been adopted by airframe manufacturers in current and future aircraft. Modern lightweight materials can provide less structural rigidity while maintaining load-carrying capacity. As structural flexibility increases, aeroelastic interactions with aerodynamic forces and moments become an increasingly important consideration in aircraft design and aerodynamic performance. Furthermore, aeroelastic interactions with flight dynamics can result in issues with vehicle stability and control. Abstract This paper describes a recent aeroelastic modeling effort for an elastically shaped aircraft concept (ESAC). The aircraft model is based on the rigid-body generic transport model (GTM) originally developed at NASA Langley Research Center. The ESAC distinguishes itself from the GTM in that it is equipped with highly flexible wing structures as a weight reduction design feature. More significantly, the wings are outfitted with a novel control effector concept called variable camber continuous trailing edge (VCCTE) flap system for active control of wing aeroelastic deflections to optimize the local angle of attack of wing sections for improved aerodynamic efficiency through cruise drag reduction and lift enhancement during take-off and landing. The VCCTE flap is a multi-functional and aerodynamically efficient device capable of achieving high lift-to-drag ratios. The flap system is comprised of three chordwise segments that form the variable camber feature of the flap and multiple spanwise segments that form a piecewise continuous trailing edge. By configuring the flap camber and trailing edge shape, drag reduction could be achieved. Moreover, some parts of the flap system can be made to have a high frequency response for roll control, gust load alleviation, and aeroservoelastic (ASE) modal suppression control. Abstract The aeroelastic model of the ESAC is based on one-dimensional structural dynamic theory that captures the aeroelastic deformation of a wing structure in a combined motion that involves flapwise bending, chordwise bending, and torsion. The model includes the effect of aircraft propulsion due to wing flexibility which causes the propulsive forces and moments to couple with the wing elastic motion. Engine mass is also accounted in the model. A fuel management model is developed to describe the wing mass change due to fuel usage in the main tank and wing tanks during cruise. Abstract The model computes both static and dynamic responses of the wing structures. The static aeroelastic deflections are used to estimate the effect of wing flexibility on induced drag and the potential drag reduction by the VCCTE flap system. A flutter analysis is conducted to estimate the flutter speed boundary. Gust load alleviation via adaptive control has been recently investigated to address flexibility of aircraft structures. A multi-objective flight control approach is presented for drag reduction control. The approach is based on an optimal control framework using a multi-objective cost function. Future studies will demonstrate the potential benefits of the approach.

Point vortex model for prediction of sound generated by a wing with flap interacting with a passing vortex.

PubMed

Manela, A; Huang, L

2013-04-01

Acoustic signature of a rigid wing, equipped with a movable downstream flap and interacting with a line vortex, is studied in a two-dimensional low-Mach number flow. The flap is attached to the airfoil via a torsion spring, and the coupled fluid-structure interaction problem is analyzed using thin-airfoil methodology and application of the emended Brown and Michael equation. It is found that incident vortex passage above the airfoil excites flap motion at the system natural frequency, amplified above all other frequencies contained in the forcing vortex. Far-field radiation is analyzed using Powell-Howe analogy, yielding the leading order dipole-type signature of the system. It is shown that direct flap motion has a negligible effect on total sound radiation. The characteristic acoustic signature of the system is dominated by vortex sound, consisting of relatively strong leading and trailing edge interactions of the airfoil with the incident vortex, together with late-time wake sound resulting from induced flap motion. In comparison with the counterpart rigid (non-flapped) configuration, it is found that the flap may act as sound amplifier or absorber, depending on the value of flap-fluid natural frequency. The study complements existing analyses examining sound radiation in static- and detached-flap configurations.

Vortical Flow Structures in the Near-Wake of a Heaving Airfoil with Passively Actuated Leading and Trailing Flaps.

NASA Astrophysics Data System (ADS)

Siala, Firas; Totpal, Alexander; Liburdy, James

2015-11-01

The flow physics of flying animals has recently received significant attention, mostly in the context of developing bio-inspired micro air vehicles and oscillating flow energy harvesters. Of particular interest is the understanding of the impact of airfoil flexibility on the flow physics. Research efforts showed that some degree of surface flexibility enhanced the strength and size of the leading edge vortex. In this study, the influence of flexibility on the near-wake dynamics and flow structures is investigated using 2D PIV measurements. The experiments are conducted in a wind tunnel at a Reynolds number of 30,000 and a range of reduced frequencies from 0.09 to 0.2. The flexibility is attained using a torsion rod forming a hinge between the flap and the main wing. Vortex flow structures are visualized using large eddy scale decomposition technique and quantified using swirling strength analysis. It is found that trailing edge flexibility increases the vortex swirling strength compared to a rigid airfoil, whereas leading edge flexibility decreases the swirling strength. Furthermore, the integral length scale determined from the autocorrelation of the velocity fluctuations is found to be approximately equal to the actual vortex size. The vortex convective velocity is shown to be independent of flexibility and oscillation frequency, and it is represented by a trimodal distribution, with peak values at 0.8, 0.95 and 1 times the free stream velocity. Oregon State University.

Numerical study on the aerodynamic characteristics of both static and flapping wing with attachments

NASA Astrophysics Data System (ADS)

Xie, Lingwang; Zhang, Xingwei; Luo, Pan; Huang, Panpan

2017-10-01

The purpose of this paper is to investigate the aerodynamic mechanism of airfoils under different icing situations which are different icing type, different icing time, and different icing position. Numerical simulation is carried out by using the finite volume method for both static and flapping airfoils, when Reynolds number is kept at 135000. The difference of aerodynamic performance between the airfoil with attachments and without attachments are be investigated by comparing the force coefficients, lift-to-drag ratios and flow field contour. The present simulations reveal that some influences of attachment are similar in the static airfoil and the flapping airfoil. Specifically, the airfoil with the attachment derived from glaze ice type causes the worse aerodynamic performance than that derived from rime ice type. The longer the icing time, the greater influence of aerodynamic performance the attachment causes. The attachments on the leading-edge have the greater influence of aerodynamic performance than other positions. Moreover, there are little differences between the static airfoil and the flapping airfoil. Compared with the static airfoil, the flapping airfoil which attachment located on the trailing edge causes a worse aerodynamic performance. Both attachments derived from rime ice type and glaze ice type all will deteriorate the aerodynamic performance of the asymmetrical airfoils. Present work provides the systematic and comprehensive study about icing blade which is conducive to the development of the wind power generation technology.

Wind-tunnel investigation of a large-scale VTOL aircraft model with wing root and wing thrust augmentors. [Ames 40 by 80 foot wind tunnel

NASA Technical Reports Server (NTRS)

Aoyagi, K.; Aiken, T. N.

1979-01-01

Tests were conducted in the Ames 40 by 80 foot wind tunnel to determine the aerodynamic characteristics of a large-scale V/STOL aircraft model with thrust augmentors. The model had a double-delta wing of aspect ratio 1.65 with augmentors located in the wing root and the wing trailing edge. The supply air for the augmentor primary nozzles was provided by the YJ-97 turbojet engine. The airflow was apportioned approximately 74 percent to the wing root augmentor and 24 percent to wing augmentor. Results were obtained at several trailing-edge flap deflections with the nozzle jet-momentum coefficients ranging from 0 to 7.9. Three-component longitudinal data are presented with the agumentor operating with and without the horizontal tail. A limited amount of six component data are also presented.

Subsonic longitudinal aerodynamic characteristics of a vectored-engine-over-wing configuration having spanwise leading-edge vortex enhancement

NASA Technical Reports Server (NTRS)

Huffman, J. K.; Fox, C. H., Jr.

1977-01-01

A configuration which integrates a close coupled canard wing combination, spanwise blowing for enhancement of the wing leading edge vortex, an engine-over-wing concept, and a wing trailing edge coanda-effect flap is studied. The data on the configuration are presented in tabular from without discussion. The investigation was conducted in the Langley 7- by 10-foot high speed tunnel at a Mach number of 0.166 through an angle-of-attack range from -2 to 22 deg. Rectangular main engine nozzles of aspect ratio 4, 6, and 8 were tested over a momentum coefficient range from 1.0 to 1.8.

Aerodynamic characteristics of a large-scale semispan model with a swept wing and an augmented jet flap with hypermixing nozzles. [Ames 40- by 80-Foot Wind Tunnel and Static Test Facility

NASA Technical Reports Server (NTRS)

Aiken, T. N.; Falarski, M. D.; Koenin, D. G.

1979-01-01

The aerodynamic characteristics of the augmentor wing concept with hypermixing primary nozzles were investigated. A large-scale semispan model in the Ames 40- by 80-Foot Wind Tunnel and Static Test Facility was used. The trailing edge, augmentor flap system occupied 65% of the span and consisted of two fixed pivot flaps. The nozzle system consisted of hypermixing, lobe primary nozzles, and BLC slot nozzles at the forward inlet, both sides and ends of the throat, and at the aft flap. The entire wing leading edge was fitted with a 10% chord slat and a blowing slot. Outboard of the flap was a blown aileron. The model was tested statically and at forward speed. Primary parameters and their ranges included angle of attack from -12 to 32 degrees, flap angles of 20, 30, 45, 60 and 70 degrees, and deflection and diffuser area ratios from 1.16 to 2.22. Thrust coefficients ranged from 0 to 2.73, while nozzle pressure ratios varied from 1.0 to 2.34. Reynolds number per foot varied from 0 to 1.4 million. Analysis of the data indicated a maximum static, gross augmentation of 1.53 at a flap angle of 45 degrees. Analysis also indicated that the configuration was an efficient powered lift device and that the net thrust was comparable with augmentor wings of similar static performance. Performance at forward speed was best at a diffuser area ratio of 1.37.

Selected topics on the active control of helicopter aeromechanical and vibration problems

NASA Technical Reports Server (NTRS)

Friedmann, Peretz P.

1994-01-01

This paper describes in a concise manner three selected topics on the active control of helicopter aeromechanical and vibration problems. The three topics are as follows: (1) the active control of helicopter air-resonance using an LQG/LTR approach; (2) simulation of higher harmonic control (HHC) applied to a four bladed hingeless helicopter rotor in forward flight; and (3) vibration suppression in forward flight on a hingeless helicopter rotor using an actively controlled, partial span, trailing edge flap, which is mounted on the blade. Only a few selected illustrative results are presented. The results obtained clearly indicate that the partial span, actively controlled flap has considerable potential for vibration reduction in helicopter rotors.

Assessment of PIV-based unsteady load determination of an airfoil with actuated flap

NASA Astrophysics Data System (ADS)

Sterenborg, J. J. H. M.; Lindeboom, R. C. J.; Simão Ferreira, C. J.; van Zuijlen, A. H.; Bijl, H.

2014-02-01

For complex experimental setups involving movable structures it is not trivial to directly measure unsteady loads. An alternative is to deduce unsteady loads indirectly from measured velocity fields using Noca's method. The ultimate aim is to use this method in future work to determine unsteady loads for fluid-structure interaction problems. The focus in this paper is first on the application and assessment of Noca's method for an airfoil with an oscillating trailing edge flap. To our best knowledge Noca's method has not been applied yet to airfoils with moving control surfaces or fluid-structure interaction problems. In addition, wind tunnel corrections for this type of unsteady flow problem are considered.

Flow Field Analysis of Fully Coupled Computations of a Flexible Wing undergoing Stall Flutter

DTIC Science & Technology

2016-01-01

unsteady aerodynamic loads due to structural displacements. In terms of actuation , most, if not all, active ∗Research Associate, Department of...flutter suppression techniques, conventional trailing edge flap actuators with a bandwidth of 10 Hz5 was used. Interestingly, the frequencies associated...influence of the flow features on the aeroelastic instability are quantified. Finally, the influence of actuation through a blowing port at 75% span is

Transonic Drag Reduction Through Trailing-Edge Blowing on the FAST-MAC Circulation Control Model

NASA Technical Reports Server (NTRS)

Chan, David T.; Jones, Gregory S.; Milholen, William E., II; Goodliff, Scott L.

2017-01-01

A third wind tunnel test of the FAST-MAC circulation control semi-span model was completed in the National Transonic Facility at the NASA Langley Research Center where the model was configured for transonic testing of the cruise configuration with 0deg flap detection to determine the potential for transonic drag reduction with the circulation control blowing. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged ap. Recent upgrades to transonic semi-span flow control testing at the NTF have demonstrated an improvement to overall data repeatability, particularly for the drag measurement, that allows for increased confidence in the data results. The static thrust generated by the blowing slot was removed from the wind-on data using force and moment balance data from wind-o thrust tares. This paper discusses the impact of the trailing-edge blowing to the transonic aerodynamics of the FAST-MAC model in the cruise configuration, where at flight Reynolds numbers, the thrust-removed corrected data showed that an overall drag reduction and increased aerodynamic efficiency was realized as a consequence of the blowing.

Shock Location Dominated Transonic Flight Loads on the Active Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Lokos, William A.; Lizotte, Andrew; Lindsley, Ned J.; Stauf, Rick

2005-01-01

During several Active Aeroelastic Wing research flights, the shadow of the over-wing shock could be observed because of natural lighting conditions. As the plane accelerated, the shock location moved aft, and as the shadow passed the aileron and trailing-edge flap hinge lines, their associated hinge moments were substantially affected. The observation of the dominant effect of shock location on aft control surface hinge moments led to this investigation. This report investigates the effect of over-wing shock location on wing loads through flight-measured data and analytical predictions. Wing-root and wing-fold bending moment and torque and leading- and trailing-edge hinge moments have been measured in flight using calibrated strain gages. These same loads have been predicted using a computational fluid dynamics code called the Euler Navier-Stokes Three Dimensional Aeroelastic Code. The computational fluid dynamics study was based on the elastically deformed shape estimated by a twist model, which in turn was derived from in-flight-measured wing deflections provided by a flight deflection measurement system. During level transonic flight, the shock location dominated the wing trailing-edge control surface hinge moments. The computational fluid dynamics analysis based on the shape provided by the flight deflection measurement system produced very similar results and substantially correlated with the measured loads data.

Use of active control systems to improve bending and rotor flapping response of a tilt rotor VTOL airplane

NASA Technical Reports Server (NTRS)

Whitaker, H. P.; Cheng, Y.

1975-01-01

The results are summarized of an analytical study of the use of active control systems for the purpose of reducing the root mean square response of wing vertical bending and rotor flapping to atmospheric turbulence for a tilt-rotor VTOL airplane. Only the wing/rotor assembly was considered so that results of a wind tunnel test program would be applicable in a subsequent phase of the research. The capabilities and limitations of simple single feedback configurations were identified, and the most promising multiloop feedback configurations were then investigated. Design parameters were selected so as to minimize either wing bending or rotor flapping response. Within the constraints imposed by practical levels of feedback gains and complexity and by considerations of safety, reduction in response due to turbulence of the order of 30 to 50 percent is predicted using the rotor longitudinal cyclic and a trailing edge wing flap as control effectors.

The use of a panel code on high lift configurations of a swept forward wing

NASA Technical Reports Server (NTRS)

Scheib, J. S.; Sandlin, D. R.

1985-01-01

A study was done on high lift configurations of a generic swept forward wing using a panel code prediction method. A survey was done of existing codes available at Ames, frow which the program VSAERO was chosen. The results of VSAERO were compared with data obtained from the Ames 7- by 10-foot wind tunnel. The results of the comparison in lift were good (within 3.5%). The comparison of the pressure coefficients was also good. The pitching moment coefficients obtained by VSAERO were not in good agreement with experiment. VSAERO's ability to predict drag is questionable and cannot be counted on for accurate trends. Further studies were done on the effects of a leading edge glove, canards, leading edge sweeps and various wing twists on spanwise loading and trim lift with encouraging results. An unsuccessful attempt was made to model spanwise blowing and boundary layer control on the trailing edge flap. The potential results of VSAERO were compared with experimental data of flap deflections with boundary layer control to check the first order effects.

A computer program to calculate the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

NASA Technical Reports Server (NTRS)

Mendenhall, M. R.; Goodwin, F. K.; Spangler, S. B.

1976-01-01

A vortex lattice lifting-surface method is used to model the wing and multiple flaps. Each lifting surface may be of arbitrary planform having camber and twist, and the multiple-slotted trailing-edge flap system may consist of up to ten flaps with different spans and deflection angles. The engine wakes model consists of a series of closely spaced vortex rings with circular or elliptic cross sections. The rings are normal to a wake centerline which is free to move vertically and laterally to accommodate the local flow field beneath the wing and flaps. The two potential flow models are used in an iterative fashion to calculate the wing-flap loading distribution including the influence of the waves from up to two turbofan engines on the semispan. The method is limited to the condition where the flow and geometry of the configurations are symmetric about the vertical plane containing the wing root chord. The calculation procedure starts with arbitrarily positioned wake centerlines and the iterative calculation continues until the total configuration loading converges within a prescribed tolerance. Program results include total configuration forces and moments, individual lifting-surface load distributions, including pressure distributions, individual flap hinge moments, and flow field calculation at arbitrary field points.

Topology optimization of pressure adaptive honeycomb for a morphing flap

NASA Astrophysics Data System (ADS)

Vos, Roelof; Scheepstra, Jan; Barrett, Ron

2011-03-01

The paper begins with a brief historical overview of pressure adaptive materials and structures. By examining avian anatomy, it is seen that pressure-adaptive structures have been used successfully in the Natural world to hold structural positions for extended periods of time and yet allow for dynamic shape changes from one flight state to the next. More modern pneumatic actuators, including FAA certified autopilot servoactuators are frequently used by aircraft around the world. Pneumatic artificial muscles (PAM) show good promise as aircraft actuators, but follow the traditional model of load concentration and distribution commonly found in aircraft. A new system is proposed which leaves distributed loads distributed and manipulates structures through a distributed actuator. By using Pressure Adaptive Honeycomb (PAH), it is shown that large structural deformations in excess of 50% strains can be achieved while maintaining full structural integrity and enabling secondary flight control mechanisms like flaps. The successful implementation of pressure-adaptive honeycomb in the trailing edge of a wing section sparked the motivation for subsequent research into the optimal topology of the pressure adaptive honeycomb within the trailing edge of a morphing flap. As an input for the optimization two known shapes are required: a desired shape in cruise configuration and a desired shape in landing configuration. In addition, the boundary conditions and load cases (including aerodynamic loads and internal pressure loads) should be specified for each condition. Finally, a set of six design variables is specified relating to the honeycomb and upper skin topology of the morphing flap. A finite-element model of the pressure-adaptive honeycomb structure is developed specifically tailored to generate fast but reliable results for a given combination of external loading, input variables, and boundary conditions. Based on two bench tests it is shown that this model correlates well to experimental results. The optimization process finds the skin and honeycomb topology that minimizes the error between the acquired shape and the desired shape in each configuration.

Experimental and computational investigation of lift-enhancing tabs on a multi-element airfoil

NASA Technical Reports Server (NTRS)

Ashby, Dale

1996-01-01

An experimental and computational investigation of the effect of lift enhancing tabs on a two-element airfoil was conducted. The objective of the study was to develop an understanding of the flow physics associated with lift enhancing tabs on a multi-element airfoil. A NACA 63(sub 2)-215 ModB airfoil with a 30 percent chord Fowler flap was tested in the NASA Ames 7 by 10 foot wind tunnel. Lift enhancing tabs of various heights were tested on both the main element and the flap for a variety of flap riggings. Computations of the flow over the two-element airfoil were performed using the two-dimensional incompressible Navier-Stokes code INS2D-UP. The computer results predict all of the trends in the experimental data quite well. When the flow over the flap upper surface is attached, tabs mounted at the main element trailing edge (cove tabs) produce very little change in lift. At high flap deflections. however, the flow over the flap is separated and cove tabs produce large increases in lift and corresponding reductions in drag by eliminating the separated flow. Cove tabs permit high flap deflection angles to be achieved and reduce the sensitivity of the airfoil lift to the size of the flap gap. Tabs attached to the flap training edge (flap tabs) are effective at increasing lift without significantly increasing drag. A combination of a cove tab and a flap tab increased the airfoil lift coefficient by 11 percent relative to the highest lift tab coefficient achieved by any baseline configuration at an angle of attack of zero percent and the maximum lift coefficient was increased by more than 3 percent. A simple analytic model based on potential flow was developed to provide a more detailed understanding of how lift enhancing tabs work. The tabs were modeled by a point vortex at the training edge. Sensitivity relationships were derived which provide a mathematical basis for explaining the effects of lift enhancing tabs on a multi-element airfoil. Results of the modeling effort indicate that the dominant effects of the tabs on the pressure distribution of each element of the airfoil can be captured with a potential flow model for cases with no flow separation.

A 3-D Computational Study of a Variable Camber Continuous Trailing Edge Flap (VCCTEF) Spanwise Segment

NASA Technical Reports Server (NTRS)

Kaul, Upender K.; Nguyen, Nhan T.

2015-01-01

Results of a computational study carried out to explore the effects of various elastomer configurations joining spanwise contiguous Variable Camber Continuous Trailing Edge Flap (VCCTEF) segments are reported here. This research is carried out as a proof-of-concept study that will seek to push the flight envelope in cruise with drag optimization as the objective. The cruise conditions can be well off design such as caused by environmental conditions, maneuvering, etc. To handle these off-design conditions, flap deflection is used so when the flap is deflected in a given direction, the aircraft angle of attack changes accordingly to maintain a given lift. The angle of attack is also a design parameter along with the flap deflection. In a previous 2D study,1 the effect of camber was investigated and the results revealed some insight into the relative merit of various camber settings of the VCCTEF. The present state of the art has not advanced sufficiently to do a full 3-D viscous analysis of the whole NASA Generic Transport Model (GTM) wing with VCCTEF deployed with elastomers. Therefore, this study seeks to explore the local effects of three contiguous flap segments on lift and drag of a model devised here to determine possible trades among various flap deflections to achieve desired lift and drag results. Although this approach is an approximation, it provides new insights into the "local" effects of the relative deflections of the contiguous spanwise flap systems and various elastomer segment configurations. The present study is a natural extension of the 2-D study to assess these local 3-D effects. Design cruise condition at 36,000 feet at free stream Mach number of 0.797 and a mean aerodynamic chord (MAC) based Reynolds number of 30.734x10(exp 6) is simulated for an angle of attack (AoA) range of 0 to 6 deg. In the previous 2-D study, the calculations revealed that the parabolic arc camber (1x2x3) and circular arc camber (VCCTEF222) offered the best L/D characteristics and minimum drag in cruise. In the present 3-D study, calculations show that for the same C(sub t), the 3-D circular arc camber wing segment produces the largest drag for a given lift, larger than either of the two 2-D configurations, as was also conjectured in the previous study. This study indicates a wing stall around 4.5 deg angle of attack.

A Mission-Adaptive Variable Camber Flap Control System to Optimize High Lift and Cruise Lift-to-Drag Ratios of Future N+3 Transport Aircraft

NASA Technical Reports Server (NTRS)

Urnes, James, Sr.; Nguyen, Nhan; Ippolito, Corey; Totah, Joseph; Trinh, Khanh; Ting, Eric

2013-01-01

Boeing and NASA are conducting a joint study program to design a wing flap system that will provide mission-adaptive lift and drag performance for future transport aircraft having light-weight, flexible wings. This Variable Camber Continuous Trailing Edge Flap (VCCTEF) system offers a lighter-weight lift control system having two performance objectives: (1) an efficient high lift capability for take-off and landing, and (2) reduction in cruise drag through control of the twist shape of the flexible wing. This control system during cruise will command varying flap settings along the span of the wing in order to establish an optimum wing twist for the current gross weight and cruise flight condition, and continue to change the wing twist as the aircraft changes gross weight and cruise conditions for each mission segment. Design weight of the flap control system is being minimized through use of light-weight shape memory alloy (SMA) actuation augmented with electric actuators. The VCCTEF program is developing better lift and drag performance of flexible wing transports with the further benefits of lighter-weight actuation and less drag using the variable camber shape of the flap.

Low-speed tests of a high-aspect-ratio, supercritical-wing transport model equipped with a high-lift flap system in the Langley 4- by 7-meter and Ames 12-foot pressure tunnels

NASA Technical Reports Server (NTRS)

Morgan, H. L., Jr.; Kjelgaard, S. O.

1983-01-01

The Ames 12-Foot Pressure Tunnel was used to determine the effects of Reynolds number on the static longitudinal aerodynamic characteristics of an advanced, high-aspect-ratio, supercritical wing transport model equipped with a full span, leading edge slat and part span, double slotted, trailing edge flaps. The model had a wing span of 7.5 ft and was tested through a free stream Reynolds number range from 1.3 to 6.0 x 10 to 6th power per foot at a Mach number of 0.20. Prior to the Ames tests, an investigation was also conducted in the Langley 4 by 7 Meter Tunnel at a Reynolds number of 1.3 x 10 to 6th power per foot with the model mounted on an Ames strut support system and on the Langley sting support system to determine strut interference corrections. The data obtained from the Langley tests were also used to compare the aerodynamic charactertistics of the rather stiff, 7.5-ft-span steel wing model tested during this investigation and the larger, and rather flexible, 12-ft-span aluminum-wing model tested during a previous investigation. During the tests in both the Langley and Ames tunnels, the model was tested with six basic wing configurations: (1) cruise; (2) climb (slats only extended); (3) 15 deg take-off flaps; (4) 30 deg take-off flaps; (5) 45 deg landing flaps; and (6) 60 deg landing flaps.

Active Flow Control Using Sweeping Jet Actuators on a Semi-Span Wing Model

NASA Technical Reports Server (NTRS)

Melton, LaTunia Pack; Koklu, Mehti

2016-01-01

Wind tunnel experiments were performed using active flow control on an unswept semispan wing model with a 30% chord trailing edge flap to aid in the selection of actuators for a planned high Reynolds number experiment. Two sweeping jet actuator sizes were investigated to determine the influence of actuator size on the active flow control system efficiency. Sweeping jet actuators with orifice sizes of 1 mm x 2 mm and 2 mm x 4 mm were selected because of the differences in actuator jet sweep angle. The parameters that were varied include actuator momentum, freestream velocity, and trailing edge flap deflection angle. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the two actuators. In addition to the wind tunnel experiments, benchtop studies of the actuators were performed to characterize the jets produced by each actuator. Benchtop investigations of the smaller actuator reveal that the jet exiting the actuator has a reduced sweep angle compared to published data for larger versions of this type of actuator. The larger actuator produces an oscillating jet that attaches to the external di?user walls at low supply pressures and produces the expected sweep angles. The AFC results using the smaller actuators show that while the actuators can control flow separation, the selected spacing of 3.3 cm may be too large due to the reduced sweep angle. In comparison, the spacing for the larger actuators, 6.6 cm, appears to be optimal for the Mach numbers investigated. Particle Image Velocimetry results are presented and show how the wall jets produced by the actuators cause the flow to attach to the flap surface.

The development of a closed-loop flight controller with panel method integration for gust alleviation using biomimetic feathers on aircraft wings

NASA Astrophysics Data System (ADS)

Blower, Christopher J.; Lee, Woody; Wickenheiser, Adam M.

2012-04-01

This paper presents the development of a biomimetic closed-loop flight controller that integrates gust alleviation and flight control into a single distributed system. Modern flight controllers predominantly rely on and respond to perturbations in the global states, resulting in rotation or displacement of the entire aircraft prior to the response. This bio-inspired gust alleviation system (GAS) employs active deflection of electromechanical feathers that react to changes in the airflow, i.e. the local states. The GAS design is a skeletal wing structure with a network of featherlike panels installed on the wing's surfaces, creating the airfoil profile and replacing the trailing-edge flaps. In this study, a dynamic model of the GAS-integrated wing is simulated to compute gust-induced disturbances. The system implements continuous adjustment to flap orientation to perform corrective responses to inbound gusts. MATLAB simulations, using a closed-loop LQR integrated with a 2D adaptive panel method, allow analysis of the morphing structure's aerodynamic data. Non-linear and linear dynamic models of the GAS are compared to a traditional single control surface baseline wing. The feedback loops synthesized rely on inertial changes in the global states; however, variations in number and location of feather actuation are compared. The bio-inspired system's distributed control effort allows the flight controller to interchange between the single and dual trailing edge flap profiles, thereby offering an improved efficiency to gust response in comparison to the traditional wing configuration. The introduction of aero-braking during continuous gusting flows offers a 25% reduction in x-velocity deviation; other flight parameters can be reduced in magnitude and deviation through control weighting optimization. Consequently, the GAS demonstrates enhancements to maneuverability and stability in turbulent intensive environments.

Loads Model Development and Analysis for the F/A-18 Active Aeroelastic Wing Airplane

NASA Technical Reports Server (NTRS)

Allen, Michael J.; Lizotte, Andrew M.; Dibley, Ryan P.; Clarke, Robert

2005-01-01

The Active Aeroelastic Wing airplane was successfully flight-tested in March 2005. During phase 1 of the two-phase program, an onboard excitation system provided independent control surface movements that were used to develop a loads model for the wing structure and wing control surfaces. The resulting loads model, which was used to develop the control laws for phase 2, is described. The loads model was developed from flight data through the use of a multiple linear regression technique. The loads model input consisted of aircraft states and control surface positions, in addition to nonlinear inputs that were calculated from flight-measured parameters. The loads model output for each wing consisted of wing-root bending moment and torque, wing-fold bending moment and torque, inboard and outboard leading-edge flap hinge moment, trailing-edge flap hinge moment, and aileron hinge moment. The development of the Active Aeroelastic Wing loads model is described, and the ability of the model to predict loads during phase 2 research maneuvers is demonstrated. Results show a good match to phase 2 flight data for all loads except inboard and outboard leading-edge flap hinge moments at certain flight conditions. The average load prediction errors for all loads at all flight conditions are 9.1 percent for maximum stick-deflection rolls, 4.4 percent for 5-g windup turns, and 7.7 percent for 4-g rolling pullouts.

Full-scale Wind-tunnel to Determine a Satisfactory Location for a Service Pitot-static Tube on a Low-wing Monoplane

NASA Technical Reports Server (NTRS)

Parsons, John F

1936-01-01

Surveys of the air flow over the upper surface of four different airfoils were made in the full-scale wind tunnel to determine a satisfactory location for a fixed Pitot-static tube on a low-wing monoplane. The selection was based on small interference errors, less than 5 percent, and on a consideration of structural and ground handling problems. The most satisfactory location on the airfoils without flaps that were investigated was 10 percent of the chord aft and 25 percent of the chord above the trailing edge of a section approximately 40 percent of the semispan inboard of the wing tip. No satisfactory location was found near the wing when the flaps were deflected.

NASA Innovation Fund 2010 Project Elastically Shaped Future Air Vehicle Concept

NASA Technical Reports Server (NTRS)

Nguyen, Nhan

2010-01-01

This report describes a study conducted in 2010 under the NASA Innovation Fund Award to develop innovative future air vehicle concepts. Aerodynamic optimization was performed to produce three different aircraft configuration concepts for low drag, namely drooped wing, inflected wing, and squashed fuselage. A novel wing shaping control concept is introduced. This concept describes a new capability of actively controlling wing shape in-flight to minimize drag. In addition, a novel flight control effector concept is developed to enable wing shaping control. This concept is called a variable camber continuous trailing edge flap that can reduce drag by as much as 50% over a conventional flap. In totality, the potential benefits of fuel savings offered by these concepts can be significant.

Aero-Mechanical Design Methodology for Subsonic Civil Transport High-Lift Systems

NASA Technical Reports Server (NTRS)

vanDam, C. P.; Shaw, S. G.; VanderKam, J. C.; Brodeur, R. R.; Rudolph, P. K. C.; Kinney, D.

2000-01-01

In today's highly competitive and economically driven commercial aviation market, the trend is to make aircraft systems simpler and to shorten their design cycle which reduces recurring, non-recurring and operating costs. One such system is the high-lift system. A methodology has been developed which merges aerodynamic data with kinematic analysis of the trailing-edge flap mechanism with minimum mechanism definition required. This methodology provides quick and accurate aerodynamic performance prediction for a given flap deployment mechanism early on in the high-lift system preliminary design stage. Sample analysis results for four different deployment mechanisms are presented as well as descriptions of the aerodynamic and mechanism data required for evaluation. Extensions to interactive design capabilities are also discussed.

System and Method for Modeling the Flow Performance Features of an Object

NASA Technical Reports Server (NTRS)

Jorgensen, Charles (Inventor); Ross, James (Inventor)

1997-01-01

The method and apparatus includes a neural network for generating a model of an object in a wind tunnel from performance data on the object. The network is trained from test input signals (e.g., leading edge flap position, trailing edge flap position, angle of attack, and other geometric configurations, and power settings) and test output signals (e.g., lift, drag, pitching moment, or other performance features). In one embodiment, the neural network training method employs a modified Levenberg-Marquardt optimization technique. The model can be generated 'real time' as wind tunnel testing proceeds. Once trained, the model is used to estimate performance features associated with the aircraft given geometric configuration and/or power setting input. The invention can also be applied in other similar static flow modeling applications in aerodynamics, hydrodynamics, fluid dynamics, and other such disciplines. For example, the static testing of cars, sails, and foils, propellers, keels, rudders, turbines, fins, and the like, in a wind tunnel, water trough, or other flowing medium.

Active Control of Separation from the Slat Shoulder of a Supercritical Airfoil

NASA Technical Reports Server (NTRS)

Pack, LaTunia G.; Schaeffler, Norman W.; Yao, Chung-Sheng; Seifert, Avi

2002-01-01

Active flow control in the form of zero-mass-flux excitation was applied at the slat shoulder of a simplified high-lift airfoil to delay flow separation. The NASA Energy Efficient Transport (EET) supercritical airfoil was equipped with a 15% chord simply hinged leading edge slat and a 25% chord simply hinged trailing edge flap. The cruise configuration data was successfully reproduced, repeating previous experiments. The effects of flap and slat deflection angles on the performance of the airfoil integral parameters were quantified. Detailed flow features were measured as well, in an attempt to identify optimal actuator placement. The measurements included: steady and unsteady model and tunnel wall pressures, wake surveys, arrays of surface hot-films, flow visualization and Particle Image Velocimetry (PIV). High frequency periodic excitation was applied to delay the occurrence of slat stall and improve the maximum lift by 10 to 15%. Low frequency amplitude modulation was used to reduce the oscillatory momentum coefficient by roughly 50% with similar aerodynamic performance.

Peak-Seeking Control For Reduced Fuel Consumption: Flight-Test Results For The Full-Scale Advanced Systems Testbed FA-18 Airplane

NASA Technical Reports Server (NTRS)

Brown, Nelson

2013-01-01

A peak-seeking control algorithm for real-time trim optimization for reduced fuel consumption has been developed by researchers at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center to address the goals of the NASA Environmentally Responsible Aviation project to reduce fuel burn and emissions. The peak-seeking control algorithm is based on a steepest-descent algorithm using a time-varying Kalman filter to estimate the gradient of a performance function of fuel flow versus control surface positions. In real-time operation, deflections of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of an F/A-18 airplane are used for optimization of fuel flow. Results from six research flights are presented herein. The optimization algorithm found a trim configuration that required approximately 3 percent less fuel flow than the baseline trim at the same flight condition. This presentation also focuses on the design of the flight experiment and the practical challenges of conducting the experiment.

Optimization of multi-element airfoils for maximum lift

NASA Technical Reports Server (NTRS)

Olsen, L. E.

1979-01-01

Two theoretical methods are presented for optimizing multi-element airfoils to obtain maximum lift. The analyses assume that the shapes of the various high lift elements are fixed. The objective of the design procedures is then to determine the optimum location and/or deflection of the leading and trailing edge devices. The first analysis determines the optimum horizontal and vertical location and the deflection of a leading edge slat. The structure of the flow field is calculated by iteratively coupling potential flow and boundary layer analysis. This design procedure does not require that flow separation effects be modeled. The second analysis determines the slat and flap deflection required to maximize the lift of a three element airfoil. This approach requires that the effects of flow separation from one or more of the airfoil elements be taken into account. The theoretical results are in good agreement with results of a wind tunnel test used to corroborate the predicted optimum slat and flap positions.

Effect of winglets on a first-generation jet transport wing. 3: Pressure and spanwise load distributions for a semispan model at Mach 0.30. [in the Langley 8 ft transonic tunnel

NASA Technical Reports Server (NTRS)

Montoya, L. C.; Jacobs, P. F.; Flechner, S. G.

1977-01-01

Pressure and spanwise load distributions on a first-generation jet transport semispan model at a Mach number of 0.30 are given for the basic wing and for configurations with an upper winglet only, upper and lower winglets, and a simple wing-tip extension. To simulate second-segment-climb lift conditions, leading- and/or trailing-edge flaps were added to some configurations.

Torsional actuation with extension-torsion composite coupling and a magnetostrictive actuator

NASA Astrophysics Data System (ADS)

Bothwell, Christopher M.; Chandra, Ramesh; Chopra, Inderjit

1995-04-01

An analytical-experimental study of using magnetostrictive actuators in conjunction with an extension-torsion coupled composite tube to actuate a rotor blade trailing-edge flap to actively control helicopter vibration is presented. Thin walled beam analysis based on Vlasov theory was used to predict the induced twist and extension in a composite tube with magnetostrictive actuation. The study achieved good correlation between theory and experiment. The Kevlar-epoxy systems showed good correlation between measured and predicted twist values.

Experimental Study of Wake / Flap Interaction Noise and the Reduction of Flap Side Edge Noise

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Stead, Daniel J.; Plassman, Gerald E.

2016-01-01

The effects of the interaction of a wake with a half-span flap on radiated noise are examined. The incident wake is generated by bars of various widths and lengths or by a simplified landing gear model. Single microphone and phased array measurements are used to isolate the effects of the wake interaction on the noise radiating from the flap side edge and flap cove regions. The effects on noise of the wake generator's geometry and relative placement with respect to the flap are assessed. Placement of the wake generators upstream of the flap side edge is shown to lead to the reduction of flap side edge noise by introducing a velocity deficit and likely altering the instabilities in the flap side edge vortex system. Significant reduction in flap side edge noise is achieved with a bar positioned directly upstream of the flap side edge. The noise reduction benefit is seen to improve with increased bar width, length and proximity to the flap edge. Positioning of the landing gear model upstream of the flap side edge also leads to decreased flap side edge noise. In addition, flap cove noise levels are significantly lower than when the landing gear is positioned upstream of the flap mid-span. The impact of the local flow velocity on the noise radiating directly from the landing gear is discussed. The effects of the landing gear side-braces on flap side edge, flap cove and landing gear noise are shown.

Aerodynamic characteristics of a large-scale hybrid upper surface blown flap model having four engines

NASA Technical Reports Server (NTRS)

Carros, R. J.; Boissevain, A. G.; Aoyagi, K.

1975-01-01

Data are presented from an investigation of the aerodynamic characteristics of large-scale wind tunnel aircraft model that utilized a hybrid-upper surface blown flap to augment lift. The hybrid concept of this investigation used a portion of the turbofan exhaust air for blowing over the trailing edge flap to provide boundary layer control. The model, tested in the Ames 40- by 80-foot Wind Tunnel, had a 27.5 deg swept wing of aspect ratio 8 and 4 turbofan engines mounted on the upper surface of the wing. The lift of the model was augmented by turbofan exhaust impingement on the wind upper-surface and flap system. Results were obtained for three flap deflections, for some variation of engine nozzle configuration and for jet thrust coefficients from 0 to 3.0. Six-component longitudinal and lateral data are presented with four engine operation and with the critical engine out. In addition, a limited number of cross-plots of the data are presented. All of the tests were made with a downwash rake installed instead of a horizontal tail. Some of these downwash data are also presented.

Numerical investigation of multi-element airfoils

NASA Technical Reports Server (NTRS)

Cummings, Russell M.

1993-01-01

The flow over multi-element airfoils with flat-plate lift-enhancing tabs was numerically investigated. Tabs ranging in height from 0.25 percent to 1.25 percent of the reference airfoil chord were studied near the trailing edge of the main-element. This two-dimensional numerical simulation employed an incompressible Navier-Stokes solver on a structured, embedded grid topology. New grid refinements were used to improve the accuracy of the solution near the overlapping grid boundaries. The effects of various tabs were studied at a constant Reynolds number on a two-element airfoil with a slotted flap. Both computed and measured results indicated that a tab in the main-element cove improved the maximum lift and lift-to-drag ratio relative to the baseline airfoil without a tab. Computed streamlines revealed that the additional turning caused by the tab may reduce the amount of separated flow on the flap. A three-element airfoil was also studied over a range of Reynolds numbers. For the optimized flap rigging, the computed and measured Reynolds number effects were similar. When the flap was moved from the optimum position, numerical results indicated that a tab may help to reoptimize the airfoil to within 1 percent of the optimum flap case.

Boeing Smart Rotor Full-scale Wind Tunnel Test Data Report

NASA Technical Reports Server (NTRS)

Kottapalli, Sesi; Hagerty, Brandon; Salazar, Denise

2016-01-01

A full-scale helicopter smart material actuated rotor technology (SMART) rotor test was conducted in the USAF National Full-Scale Aerodynamics Complex 40- by 80-Foot Wind Tunnel at NASA Ames. The SMART rotor system is a five-bladed MD 902 bearingless rotor with active trailing-edge flaps. The flaps are actuated using piezoelectric actuators. Rotor performance, structural loads, and acoustic data were obtained over a wide range of rotor shaft angles of attack, thrust, and airspeeds. The primary test objective was to acquire unique validation data for the high-performance computing analyses developed under the Defense Advanced Research Project Agency (DARPA) Helicopter Quieting Program (HQP). Other research objectives included quantifying the ability of the on-blade flaps to achieve vibration reduction, rotor smoothing, and performance improvements. This data set of rotor performance and structural loads can be used for analytical and experimental comparison studies with other full-scale rotor systems and for analytical validation of computer simulation models. The purpose of this final data report is to document a comprehensive, highquality data set that includes only data points where the flap was actively controlled and each of the five flaps behaved in a similar manner.

Aerodynamics Characteristics of Multi-Element Airfoils at -90 Degrees Incidence

NASA Technical Reports Server (NTRS)

Stremel, Paul M.; Schmitz, Fredric H. (Technical Monitor)

1994-01-01

A developed method has been applied to calculate accurately the viscous flow about airfoils normal to the free-stream flow. This method has special application to the analysis of tilt rotor aircraft in the evaluation of download. In particular, the flow about an XV-15 airfoil with and without deflected leading and trailing edge flaps at -90 degrees incidence is evaluated. The multi-element aspect of the method provides for the evaluation of slotted flap configurations which may lead to decreased drag. The method solves for turbulent flow at flight Reynolds numbers. The flow about the XV-15 airfoil with and without flap deflections has been calculated and compared with experimental data at a Reynolds number of one million. The comparison between the calculated and measured pressure distributions are very good, thereby, verifying the method. The aerodynamic evaluation of multielement airfoils will be conducted to determine airfoil/flap configurations for reduced airfoil drag. Comparisons between the calculated lift, drag and pitching moment on the airfoil and the airfoil surface pressure will also be presented.

Effects of leading-edge devices on the low-speed aerodynamic characteristics of a highly-swept arrow-wing

NASA Technical Reports Server (NTRS)

Scott, S. J.; Nicks, O. W.; Imbrie, P. K.

1985-01-01

An investigation was conducted in the Texas A&M University 7 by 10 foot Low Speed Wind Tunnel to provide a direct comparison of the effect of several leading edge devices on the aerodynamic performance of a highly swept wing configuration. Analysis of the data indicates that for the configuration with undeflected leading edges, vortex separation first occurs on the outboard wing panel for angles of attack of approximately 2, and wing apex vorticies become apparent for alpha or = 4 deg. However, the occurrence of the leading edge vortex flow may be postponed with leading edge devices. Of the devices considered, the most promising were a simple leading edge deflection of 30 deg and a leading edge slat system. The trailing edge flap effectiveness was found to be essentially the same for the configuration employing either of these more promising leading edge devices. Analysis of the lateral directional data showed that for all of the concepts considered, deflecting leading edge downward in an attempt to postpone leading edge vortex flows, has the favorable effect of reducing the effective dihedral.

Atmospheric Probe Model: Construction and Wind Tunnel Tests

NASA Technical Reports Server (NTRS)

Vogel, Jerald M.

1998-01-01

The material contained in this document represents a summary of the results of a low speed wind tunnel test program to determine the performance of an atmospheric probe at low speed. The probe configuration tested consists of a 2/3 scale model constructed from a combination of hard maple wood and aluminum stock. The model design includes approximately 130 surface static pressure taps. Additional hardware incorporated in the baseline model provides a mechanism for simulating external and internal trailing edge split flaps for probe flow control. Test matrix parameters include probe side slip angle, external/internal split flap deflection angle, and trip strip applications. Test output database includes surface pressure distributions on both inner and outer annular wings and probe center line velocity distributions from forward probe to aft probe locations.

Effect of simulated forward airspeed on small-scale-model externally blown flap noise

NASA Technical Reports Server (NTRS)

Goodykoontz, J. H.; Dorsch, R. G.; Olsen, W. A.

1976-01-01

Noise tests were conducted on a small-scale model of an externally blown flap lift augmentation system. The nozzle/wing model was subjected to external flow that simulated takeoff and landing flight velocities by placing it in a 33-centimeter-diameter free jet. The results showed that external flow attenuated the noise associated with the various configurations tested. The amount of attenuation depended on flap setting. More attenuation occurred with a trailing-flap setting of 20 deg than with one of 60 deg. Noise varied with relative velocity as a function of the trailing-flap setting and the angle from the nozzle inlet.

Flight Wing Surface Pressure and Boundary-Layer Data Report from the F-111 Smooth Variable-Camber Supercritical Mission Adaptive Wing

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.

1997-01-01

Flight tests were conducted using the advanced fighter technology integration F-111 (AFTI/F-111) aircraft modified with a variable-sweep supercritical mission adaptive wing (MAW). The MAW leading- and trailing-edge variable-camber surfaces were deflected in flight to provide a near-ideal wing camber shape for the flight condition. The MAW features smooth, flexible upper surfaces and fully enclosed lower surfaces, which distinguishes it from conventional flaps that have discontinuous surfaces and exposed or semi-exposed mechanisms. Upper and lower surface wing pressure distributions were measured along four streamwise rows on the right wing for cruise, maneuvering, and landing configurations. Boundary-layer measurements were obtained near the trailing edge for one of the rows. Cruise and maneuvering wing leading-edge sweeps were 26 deg for Mach numbers less than 1 and 45 deg or 58 deg for Mach numbers greater than 1. The landing wing sweep was 9 deg or 16 deg. Mach numbers ranged from 0.27 to 1.41, angles of attack from 2 deg to 13 deg, and Reynolds number per unit foot from 1.4 x 10(exp 6) to 6.5 x 10(exp 6). Leading-edge cambers ranged from O deg to 20 deg down, and trailing-edge cambers ranged from 1 deg up to 19 deg down. Wing deflection data for a Mach number of 0.85 are shown for three cambers. Wing pressure and boundary-layer data are given. Selected data comparisons are shown. Measured wing coordinates are given for three streamwise semispan locations for cruise camber and one spanwise location for maneuver camber.

Elastomeric Structural Attachment Concepts for Aircraft Flap Noise Reduction - Challenges and Approaches to Hyperelastic Structural Modeling and Analysis

NASA Technical Reports Server (NTRS)

Sreekantamurthy, Thammaiah; Turner, Travis L.; Moore, James B.; Su, Ji

2014-01-01

Airframe noise is a significant part of the overall noise of transport aircraft during the approach and landing phases of flight. Airframe noise reduction is currently emphasized under the Environmentally Responsible Aviation (ERA) and Fixed Wing (FW) Project goals of NASA. A promising concept for trailing-edge-flap noise reduction is a flexible structural element or link that connects the side edges of the deployable flap to the adjacent main-wing structure. The proposed solution is distinguished by minimization of the span-wise extent of the structural link, thereby minimizing the aerodynamic load on the link structure at the expense of increased deformation requirement. Development of such a flexible structural link necessitated application of hyperelastic materials, atypical structural configurations and novel interface hardware. The resulting highly-deformable structural concept was termed the FLEXible Side Edge Link (FLEXSEL) concept. Prediction of atypical elastomeric deformation responses from detailed structural analysis was essential for evaluating feasible concepts that met the design constraints. The focus of this paper is to describe the many challenges encountered with hyperelastic finite element modeling and the nonlinear structural analysis of evolving FLEXSEL concepts. Detailed herein is the nonlinear analysis of FLEXSEL concepts that emerged during the project which include solid-section, foamcore, hollow, extended-span and pre-stressed concepts. Coupon-level analysis performed on elastomeric interface joints, which form a part of the FLEXSEL topology development, are also presented.

Visualization of the separation and subsequent transition near the leading edge of airfoils

NASA Technical Reports Server (NTRS)

Arena, A. V.; Mueller, T. J.

1978-01-01

A visual study was performed using the low speed smoke wind tunnels with the objective of obtaining a better understanding of the structure of leading edge separation bubbles on airfoils. The location of separation, transition and reattachment for a cylindrical nose constant-thickness airfoil model were obtained from smoke photographs and surface oil flow techniques. These data, together with static pressure distributions along the leading edge and upper surface of the model, produced the influence of Reynolds number, angle of attack, and trailing edge flap angle on the size and characteristics of the bubble. Additional visual insight into the unsteady nature of the separation bubble was provided by high speed 16 mm movies. The 8 mm color movies taken of the surface oil flow supported the findings of the high speed movies and clearly showed the formation of a scalloped spanwise separation line at the higher Reynolds number.

Mechanisms and actuators for rotorcraft blade morphing

NASA Astrophysics Data System (ADS)

Vocke, Robert D., III

The idea of improved fight performance through changes in the control surfaces dates back to the advent of aviation with the Wright brothers' pioneering work on "wing warping," but it was not until the recent progress in material and actuator development that such control surfaces seemed practical for modern aircraft. This has opened the door to a new class of aircraft that have the ability to change shape or morph, which are being investigated due to the potential to have a single platform serve multiple mission objectives, as well as improve performance characteristics. While the majority of existing research for morphing aircraft has focused on fixedwing aircraft, rotary-wing aircraft have begun to receive more attention. The purpose of this body of work is to investigate the current state of morphing actuation technology for rotorcraft and improve upon it. Specifically, this work looks at two types of morphing: Pneumatic Artificial Muscle (PAM) actuated trailing edge flaps and conformal variable diameter morphing. First, active camber changes through the use of PAM powered trailing edge flaps were investigated due to the potential for reductions in power requirements and vibration/noise levels. A PAM based antagonistic actuation system was developed utilizing a novel combination of mechanism geometry and PAM bias contraction optimization to overcome the natural extension stiffening characteristics of PAMs. In open-loop bench-top testing against a "worst-case" constant torsional loading, the system demonstrated actuation authority suitable for both primary control and vibration/noise reduction. Additionally, closed-loop test data indicated that the system was capable of tracking complex waveforms consistent with those needed for rotorcraft control. This system demonstrated performance on-par with the state of the art pneumatic trailing edge flap actuators, yet with a much smaller footprint and impact on the rotor-blade. The second morphing system developed in this work is a conformal variable diameter rotor system suitable for implementation on a modern tilt-rotor aircraft, which can reduce power requirements in both cruise and hover configurations. An initial prototype variable span airfoil was constructed using a silicone elastomer matrix composite skin and a plastic rapid prototyped morphing substructure. Benchtop and wind tunnel tests verified the ability of this system to increase active wing area by 100%. The prototype technology was then matured for use in the harsh rotor blade environment, with a much stiffer polyurethane skin and a titanium substructure. Coupon testing verified the efficacy of this approach, and a final conceptual design was completed using the stiffness-tuning characteristics of the morphing substructure to create a self-actuating morphing blade tip.

The leading-edge vortex of swift wing-shaped delta wings

NASA Astrophysics Data System (ADS)

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

The leading-edge vortex of swift wing-shaped delta wings

PubMed Central

Muir, Rowan Eveline; Arredondo-Galeana, Abel

2017-01-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus. The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing. PMID:28878968

The leading-edge vortex of swift wing-shaped delta wings.

PubMed

Muir, Rowan Eveline; Arredondo-Galeana, Abel; Viola, Ignazio Maria

2017-08-01

Recent investigations on the aerodynamics of natural fliers have illuminated the significance of the leading-edge vortex (LEV) for lift generation in a variety of flight conditions. A well-documented example of an LEV is that generated by aircraft with highly swept, delta-shaped wings. While the wing aerodynamics of a manoeuvring aircraft, a bird gliding and a bird in flapping flight vary significantly, it is believed that this existing knowledge can serve to add understanding to the complex aerodynamics of natural fliers. In this investigation, a model non-slender delta-shaped wing with a sharp leading edge is tested at low Reynolds number, along with a delta wing of the same design, but with a modified trailing edge inspired by the wing of a common swift Apus apus . The effect of the tapering swift wing on LEV development and stability is compared with the flow structure over the unmodified delta wing model through particle image velocimetry. For the first time, a leading-edge vortex system consisting of a dual or triple LEV is recorded on a swift wing-shaped delta wing, where such a system is found across all tested conditions. It is shown that the spanwise location of LEV breakdown is governed by the local chord rather than Reynolds number or angle of attack. These findings suggest that the trailing-edge geometry of the swift wing alone does not prevent the common swift from generating an LEV system comparable with that of a delta-shaped wing.

Non-linear controls influence functions in an aircraft dynamics simulator

NASA Technical Reports Server (NTRS)

Guerreiro, Nelson M.; Hubbard, James E., Jr.; Motter, Mark A.

2006-01-01

In the development and testing of novel structural and controls concepts, such as morphing aircraft wings, appropriate models are needed for proper system characterization. In most instances, available system models do not provide the required additional degrees of freedom for morphing structures but may be modified to some extent to achieve a compatible system. The objective of this study is to apply wind tunnel data collected for an Unmanned Air Vehicle (UAV), that implements trailing edge morphing, to create a non-linear dynamics simulator, using well defined rigid body equations of motion, where the aircraft stability derivatives change with control deflection. An analysis of this wind tunnel data, using data extraction algorithms, was performed to determine the reference aerodynamic force and moment coefficients for the aircraft. Further, non-linear influence functions were obtained for each of the aircraft s control surfaces, including the sixteen trailing edge flap segments. These non-linear controls influence functions are applied to the aircraft dynamics to produce deflection-dependent aircraft stability derivatives in a non-linear dynamics simulator. Time domain analysis of the aircraft motion, trajectory, and state histories can be performed using these nonlinear dynamics and may be visualized using a 3-dimensional aircraft model. Linear system models can be extracted to facilitate frequency domain analysis of the system and for control law development. The results of this study are useful in similar projects where trailing edge morphing is employed and will be instrumental in the University of Maryland s continuing study of active wing load control.

Peak-Seeking Optimization of Trim for Reduced Fuel Consumption: Flight-Test Results

NASA Technical Reports Server (NTRS)

Brown, Nelson Andrew; Schaefer, Jacob Robert

2013-01-01

A peak-seeking control algorithm for real-time trim optimization for reduced fuel consumption has been developed by researchers at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center to address the goals of the NASA Environmentally Responsible Aviation project to reduce fuel burn and emissions. The peak-seeking control algorithm is based on a steepest-descent algorithm using a time-varying Kalman filter to estimate the gradient of a performance function of fuel flow versus control surface positions. In real-time operation, deflections of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of an F/A-18 airplane (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) are used for optimization of fuel flow. Results from six research flights are presented herein. The optimization algorithm found a trim configuration that required approximately 3 percent less fuel flow than the baseline trim at the same flight condition. The algorithm consistently rediscovered the solution from several initial conditions. These results show that the algorithm has good performance in a relevant environment.

Code Calibration Applied to the TCA High-Lift Model in the 14 x 22 Wind Tunnel (Simulation With and Without Model Post-Mount)

NASA Technical Reports Server (NTRS)

Lessard, Wendy B.

1999-01-01

The objective of this study is to calibrate a Navier-Stokes code for the TCA (30/10) baseline configuration (partial span leading edge flaps were deflected at 30 degs. and all the trailing edge flaps were deflected at 10 degs). The computational results for several angles of attack are compared with experimental force, moments, and surface pressures. The code used in this study is CFL3D; mesh sequencing and multi-grid were used to full advantage to accelerate convergence. A multi-grid approach was used similar to that used for the Reference H configuration allowing point-to-point matching across all the trailingedge block interfaces. From past experiences with the Reference H (ie, good force, moment, and pressure comparisons were obtained), it was assumed that the mounting system would produce small effects; hence, it was not initially modeled. However, comparisons of lower surface pressures indicated the post mount significantly influenced the lower surface pressures, so the post geometry was inserted into the existing grid using Chimera (overset grids).

Peak-Seeking Optimization of Trim for Reduced Fuel Consumption: Flight-test Results

NASA Technical Reports Server (NTRS)

Brown, Nelson Andrew; Schaefer, Jacob Robert

2013-01-01

A peak-seeking control algorithm for real-time trim optimization for reduced fuel consumption has been developed by researchers at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center to address the goals of the NASA Environmentally Responsible Aviation project to reduce fuel burn and emissions. The peak-seeking control algorithm is based on a steepest-descent algorithm using a time-varying Kalman filter to estimate the gradient of a performance function of fuel flow versus control surface positions. In real-time operation, deflections of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of an F/A-18 airplane (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) are used for optimization of fuel flow. Results from six research flights are presented herein. The optimization algorithm found a trim configuration that required approximately 3 percent less fuel flow than the baseline trim at the same flight condition. The algorithm consistently rediscovered the solution from several initial conditions. These results show that the algorithm has good performance in a relevant environment.

Peak Seeking Control for Reduced Fuel Consumption with Preliminary Flight Test Results

NASA Technical Reports Server (NTRS)

Brown, Nelson

2012-01-01

The Environmentally Responsible Aviation project seeks to accomplish the simultaneous reduction of fuel burn, noise, and emissions. A project at NASA Dryden Flight Research Center is contributing to ERAs goals by exploring the practical application of real-time trim configuration optimization for enhanced performance and reduced fuel consumption. This peak-seeking control approach is based on Newton-Raphson algorithm using a time-varying Kalman filter to estimate the gradient of the performance function. In real-time operation, deflection of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of a modified F-18 are directly optimized, and the horizontal stabilators and angle of attack are indirectly optimized. Preliminary results from three research flights are presented herein. The optimization system found a trim configuration that required approximately 3.5% less fuel flow than the baseline trim at the given flight condition. The algorithm consistently rediscovered the solution from several initial conditions. These preliminary results show the algorithm has good performance and is expected to show similar results at other flight conditions and aircraft configurations.

Semi-empirical airframe noise prediction model

NASA Technical Reports Server (NTRS)

Hersh, A. S.; Putnam, T. W.; Lasagna, P. L.; Burcham, F. W., Jr.

1976-01-01

A semi-empirical maximum overall sound pressure level (OASPL) airframe noise model was derived. The noise radiated from aircraft wings and flaps was modeled by using the trailing-edge diffracted quadrupole sound theory derived by Ffowcs Williams and Hall. The noise radiated from the landing gear was modeled by using the acoustic dipole sound theory derived by Curle. The model was successfully correlated with maximum OASPL flyover noise measurements obtained at the NASA Dryden Flight Research Center for three jet aircraft - the Lockheed JetStar, the Convair 990, and the Boeing 747 aircraft.

Experimental study of the separating confluent boundary-layer. Volume 2: Experimental data

NASA Technical Reports Server (NTRS)

Braden, J. A.; Whipkey, R. R.; Jones, G. S.; Lilley, D. E.

1983-01-01

An experimental low speed study of the separating confluent boundary layer on a NASA GAW-1 high lift airfoil is described. The airfoil was tested in a variety of high lift configurations comprised of leading edge slat and trailing edge flap combinations. The primary test instrumentation was a two dimensional laser velocimeter (LV) system operating in a backscatter mode. Surface pressures and corresponding LV derived boundary layer profiles are given in terms of velocity components, turbulence intensities and Reynolds shear stresses as characterizing confluent boundary layer behavior up to and beyond stall. LV derived profiles and associated boundary layer parameters and those obtained from more conventional instrumentation such as pitot static transverse, Preston tube measurements and hot-wire surveys are compared.

Vortex leading edge flap assembly for supersonic airplanes

NASA Technical Reports Server (NTRS)

Rudolph, Peter K. C. (Inventor)

1997-01-01

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.

Smart wing wind tunnel test results

NASA Astrophysics Data System (ADS)

Scherer, Lewis B.; Martin, Christopher A.; Appa, Kari; Kudva, Jayanth N.; West, Mark N.

1997-05-01

The use of smart materials technologies can provide unique capabilities in improving aircraft aerodynamic performance. Northrop Grumman built and tested a 16% scale semi-span wind tunnel model of the F/A-18 E/F for the on-going DARPA/WL Smart Materials and Structures-Smart Wing Program. Aerodynamic performance gains to be validated included increase in the lift to drag ratio, increased pitching moment (Cm), increased rolling moment (Cl) and improved pressure distribution. These performance gains were obtained using hingeless, contoured trailing edge control surfaces with embedded shape memory alloy (SMA) wires and spanwise wing twist via a SMA torque tube and are compared to a conventional wind tunnel model with hinged control surfaces. This paper presents an overview of the results from the first wind tunnel test performed at the NASA Langley's 16 ft Transonic Dynamic Tunnel. Among the benefits demonstrated are 8 - 12% increase in rolling moment due to wing twist, a 10 - 15% increase in rolling moment due to contoured aileron, and approximately 8% increase in lift due to contoured flap, and improved pressure distribution due to trailing edge control surface contouring.

A Novel SMA-based Concept for Airfoil Structural Morphing

NASA Astrophysics Data System (ADS)

Barbarino, S.; Pecora, R.; Lecce, L.; Concilio, A.; Ameduri, S.; Calvi, E.

2009-08-01

The adaptive structures concept is of great interest in the aerospace field because of the several benefits which can be accomplished in the fields including noise reduction, load alleviation, weight reduction, etc., at a level in which they can be considered as compulsory in the design of future aircraft. Improvements in terms of the aerodynamic efficiency, aeroelastic behavior, stability, and manoeuvrability performance have already been proved through many international studies in the past. In the family of the Smart Materials, Shape Memory Alloys (SMA) seem to be a suitable solution for many static applications. Their high structural integrability in conjunction with actuation capabilities and a favorable performance per weight ratio, allows the development of original architectures. In this study, a morphing wing trailing edge concept is presented; morphing ability was introduced with the aim of replacing a conventional flap device. A compliant rib structure was designed, based on SMA actuators exhibiting structural potential (bearing external aerodynamic loads). Numerical results, achieved through a FE approach, are presented in terms of trailing edge induced displacement and morphed shape.

Fish robotics and hydrodynamics

NASA Astrophysics Data System (ADS)

Lauder, George

2010-11-01

Studying the fluid dynamics of locomotion in freely-swimming fishes is challenging due to difficulties in controlling fish behavior. To provide better control over fish-like propulsive systems we have constructed a variety of fish-like robotic test platforms that range from highly biomimetic models of fins, to simple physical models of body movements during aquatic locomotion. First, we have constructed a series of biorobotic models of fish pectoral fins with 5 fin rays that allow detailed study of fin motion, forces, and fluid dynamics associated with fin-based locomotion. We find that by tuning fin ray stiffness and the imposed motion program we can produce thrust both on the fin outstroke and instroke. Second, we are using a robotic flapping foil system to study the self-propulsion of flexible plastic foils of varying stiffness, length, and trailing edge shape as a means of investigating the fluid dynamic effect of simple changes in the properties of undulating bodies moving through water. We find unexpected non-linear stiffness-dependent effects of changing foil length on self-propelled speed, and as well as significant effects of trailing edge shape on foil swimming speed.

Reduction of structural loads using maneuver load control on the Advanced Fighter Technology Integration (AFTI)/F-111 mission adaptive wing

NASA Technical Reports Server (NTRS)

Thornton, Stephen V.

1993-01-01

A transonic fighter-bomber aircraft, having a swept supercritical wing with smooth variable-camber flaps was fitted with a maneuver load control (MLC) system that implements a technique to reduce the inboard bending moments in the wing by shifting the spanwise load distribution inboard as load factor increases. The technique modifies the spanwise camber distribution by automatically commanding flap position as a function of flap position, true airspeed, Mach number, dynamic pressure, normal acceleration, and wing sweep position. Flight test structural loads data were obtained for loads in both the wing box and the wing root. Data from uniformly deflected flaps were compared with data from flaps in the MLC configuration where the outboard segment of three flap segments was deflected downward less than the two inboard segments. The changes in the shear loads in the forward wing spar and at the roots of the stabilators also are presented. The camber control system automatically reconfigures the flaps through varied flight conditions. Configurations having both moderate and full trailing-edge flap deflection were tested. Flight test data were collected at Mach numbers of 0.6, 0.7, 0.8, and 0.9 and dynamic pressures of 300, 450, 600, and 800 lb/sq ft. The Reynolds numbers for these flight conditions ranged from 26 x 10(exp 6) to 54 x 10(exp 6) at the mean aerodynamic chord. Load factor increases of up to 1.0 g achieved with no increase in wing root bending moment with the MLC flap configuration.

Active Flow Control at Low Reynolds Numbers on a NACA 0015 Airfoil

NASA Technical Reports Server (NTRS)

Melton, LaTunia Pack; Hannon, Judith; Yao, Chung-Sheng; Harris, Jerome

2008-01-01

Results from a low Reynolds number wind tunnel experiment on a NACA 0015 airfoil with a 30% chord trailing edge flap tested at deflection angles of 0, 20, and 40 are presented and discussed. Zero net mass flux periodic excitation was applied at the ap shoulder to control flow separation for flap deflections larger than 0. The primary objective of the experiment was to compare force and moment data obtained from integrating surface pressures to data obtained from a 5-component strain-gage balance in preparation for additional three-dimensional testing of the model. To achieve this objective, active flow control is applied at an angle of attack of 6 where published results indicate that oscillatory momentum coefficients exceeding 1% are required to delay separation. Periodic excitation with an oscillatory momentum coefficient of 1.5% and a reduced frequency of 0.71 caused a significant delay of separation on the airfoil with a flap deflection of 20. Higher momentum coefficients at the same reduced frequency were required to achieve a similar level of flow attachment on the airfoil with a flap deflection of 40. There was a favorable comparison between the balance and integrated pressure force and moment results.

Investigation of leading-edge flap performance on delta and double-delta wings at supersonic speeds

NASA Technical Reports Server (NTRS)

Covell, Peter F.; Wood, Richard M.; Miller, David S.

1987-01-01

An investigation of the aerodynamic performance of leading-edge flaps on three clipped delta and three clipped double-delta wing planforms with aspect ratios of 1.75, 2.11, and 2.50 was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.90, and 2.16. A primary set of fullspan leading-edge flaps with similar root and tip chords were investigated on each wing, and several alternate flap planforms were investigated on the aspect-ratio-1.75 wings. All leading-edge flap geometries were effective in reducing the drag at lifting conditions over the range of wing aspect ratios and Mach numbers tested. Application of a primary flap resulted in better flap performance with the double-delta planform than with the delta planform. The primary flap geometry generally yielded better performance than the alternate flap geometries tested. Trim drag due to flap-induced pitching moments was found to reduce the leading-edge flap performance more for the delta planform than for the double-delta planform. Flow-visualization techniques showed that leading-edge flap deflection reduces crossflow shock-induced separation effects. Finally, it was found that modified linear theory consistently predicts only the effects of leading-edge flap deflection as related to pitching moment and lift trends.

Deformation Measurements of Smart Aerodynamic Surfaces

NASA Technical Reports Server (NTRS)

Fleming, Gary A.; Burner, Alpheus

2005-01-01

Video Model Deformation (VMD) and Projection Moire Interferometry (PMI) were used to acquire wind tunnel model deformation measurements of the Northrop Grumman-built Smart Wing tested in the NASA Langley Transonic Dynamics Tunnel. The F18-E/F planform Smart Wing was outfitted with embedded shape memory alloys to actuate a seamless trailing edge aileron and flap, and an embedded torque tube to generate wing twist. The VMD system was used to obtain highly accurate deformation measurements at three spanwise locations along the main body of the wing, and at spanwise locations on the flap and aileron. The PMI system was used to obtain full-field wing shape and deformation measurements over the entire wing lower surface. Although less accurate than the VMD system, the PMI system revealed deformations occurring between VMD target rows indistinguishable by VMD. This paper presents the VMD and PMI techniques and discusses their application in the Smart Wing test.

Aerodynamic characteristics of a propulsive wing-canard concept at STOL speeds

NASA Technical Reports Server (NTRS)

Stewart, V. R.

1985-01-01

A full span model of a wing/canard concept representing a fighter configuration has been tested at STOL conditions in the NASA Langley 4 x 7 meter tunnel. The results of this test are presented, and comparisons are made to previous data of the same configuration tested as a semispan model. The potential of the propulsive wing/canard to develop very high lift coefficients was investigated with several nozzle spans (nozzle aspect ratios). Although longitudinal trim was not accomplished with the blowing distributions and configurations tested, the propulsive wing/canard appears to offer an approach to managing the large negative pitching moments associated with trailing edge flap blowing. Also presented are data showing the effects of large flap deflections and relative wing/canard positions. Presented in the appendix to the report are limited lateral-directional and ground effects data, as well as wing downwash measurements.

Sweeping Jet Optimization Studies

NASA Technical Reports Server (NTRS)

Melton, LaTunia Pack; Koklu, Mehti; Andino, Marlyn; Lin, John C.; Edelman, Louis

2016-01-01

Progress on experimental efforts to optimize sweeping jet actuators for active flow control (AFC) applications with large adverse pressure gradients is reported. Three sweeping jet actuator configurations, with the same orifice size but di?erent internal geometries, were installed on the flap shoulder of an unswept, NACA 0015 semi-span wing to investigate how the output produced by a sweeping jet interacts with the separated flow and the mechanisms by which the flow separation is controlled. For this experiment, the flow separation was generated by deflecting the wing's 30% chord trailing edge flap to produce an adverse pressure gradient. Steady and unsteady pressure data, Particle Image Velocimetry data, and force and moment data were acquired to assess the performance of the three actuator configurations. The actuator with the largest jet deflection angle, at the pressure ratios investigated, was the most efficient at controlling flow separation on the flap of the model. Oil flow visualization studies revealed that the flow field controlled by the sweeping jets was more three-dimensional than expected. The results presented also show that the actuator spacing was appropriate for the pressure ratios examined.

The Effect of Nozzle Trailing Edge Thickness on Jet Noise

NASA Technical Reports Server (NTRS)

Henderson, Brenda; Kinzie, Kevin; Haskin, Henry

2004-01-01

The effect of nozzle trailing edge thickness on broadband acoustic radiation and the production of tones is investigated for coannular nozzles. Experiments were performed for a core nozzle trailing edge thickness between 0.38 mm and 3.17 mm. The on-set of discrete tones was found to be predominantly affected by the velocity ratio, the ratio of the fan velocity to the core velocity, although some dependency on trailing edge thickness was also noted. For a core nozzle trailing edge thickness greater than or equal to 0.89 mm, tones were produced for velocity ratios between 0.91 and 1.61. For a constant nozzle trailing edge thickness, the frequency varied almost linearly with the core velocity. The Strouhal number based on the core velocity changed with nozzle trailing edge thickness and varied between 0.16 and 0.2 for the core nozzles used in the experiments. Increases in broadband noise with increasing trailing edge thickness were observed for tone producing and non-tone producing conditions. A variable thickness trailing edge (crenellated) nozzle resulted in no tonal production and a reduction of the broadband trailing edge noise relative to that of the corresponding constant thickness trailing edge.

Peak-Seeking Optimization of Spanwise Lift Distribution for Wings in Formation Flight

NASA Technical Reports Server (NTRS)

Hanson, Curtis E.; Ryan, Jack

2012-01-01

A method is presented for the in-flight optimization of the lift distribution across the wing for minimum drag of an aircraft in formation flight. The usual elliptical distribution that is optimal for a given wing with a given span is no longer optimal for the trailing wing in a formation due to the asymmetric nature of the encountered flow field. Control surfaces along the trailing edge of the wing can be configured to obtain a non-elliptical profile that is more optimal in terms of minimum combined induced and profile drag. Due to the difficult-to-predict nature of formation flight aerodynamics, a Newton-Raphson peak-seeking controller is used to identify in real time the best aileron and flap deployment scheme for minimum total drag. Simulation results show that the peak-seeking controller correctly identifies an optimal trim configuration that provides additional drag savings above those achieved with conventional anti-symmetric aileron trim.

Fiber Optic Control System Integration program: for optical flight control system development

NASA Astrophysics Data System (ADS)

Weaver, Thomas L.; Seal, Daniel W.

1994-10-01

Hardware and software were developed for optical feedback links in the flight control system of an F/A-18 aircraft. Developments included passive optical sensors and optoelectronics to operate the sensors. Sensors with different methods of operation were obtained from different manufacturers and integrated with common optoelectronics. The sensors were the following: Air Data Temperature; Air Data Pressure; and Leading Edge Flap, Nose Wheel Steering, Trailing Edge Flap, Pitch Stick, Rudder, Rudder Pedal, Stabilator, and Engine Power Lever Control Position. The sensors were built for a variety of aircraft locations and harsh environments. The sensors and optoelectronics were as similar as practical to a production system. The integrated system was installed by NASA for flight testing. Wavelength Division Multiplexing proved successful as a system design philosophy. Some sensors appeared to be better choices for aircraft applications than others, with digital sensors generally being better than analog sensors, and rotary sensors generally being better than linear sensors. The most successful sensor approaches were selected for use in a follow-on program in which the sensors will not just be flown on the aircraft and their performance recorded; but, the optical sensors will be used in closing flight control loops.

WINGDES2 - WING DESIGN AND ANALYSIS CODE

NASA Technical Reports Server (NTRS)

Carlson, H. W.

1994-01-01

This program provides a wing design algorithm based on modified linear theory which takes into account the effects of attainable leading-edge thrust. A primary objective of the WINGDES2 approach is the generation of a camber surface as mild as possible to produce drag levels comparable to those attainable with full theoretical leading-edge thrust. WINGDES2 provides both an analysis and a design capability and is applicable to both subsonic and supersonic flow. The optimization can be carried out for designated wing portions such as leading and trailing edge areas for the design of mission-adaptive surfaces, or for an entire planform such as a supersonic transport wing. This program replaces an earlier wing design code, LAR-13315, designated WINGDES. WINGDES2 incorporates modifications to improve numerical accuracy and provides additional capabilities. A means of accounting for the presence of interference pressure fields from airplane components other than the wing and a direct process for selection of flap surfaces to approach the performance levels of the optimized wing surfaces are included. An increased storage capacity allows better numerical representation of those configurations that have small chord leading-edge or trailing-edge design areas. WINGDES2 determines an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. The objective of the design is the recovery of unrealized theoretical leading-edge thrust of the input flat surface by shaping of the design surface to create a distributed thrust and thus minimize drag. The input consists of airfoil section thickness data, leading and trailing edge planform geometry, and operational parameters such as Mach number, Reynolds number, and design lift coefficient. Output includes optimized camber surface ordinates, pressure coefficient distributions, and theoretical aerodynamic characteristics. WINGDES2 is written in FORTRAN V for batch execution and has been implemented on a CDC CYBER computer operating under NOS 2.7.1 with a central memory requirement of approximately 344K (octal) of 60 bit words. This program was developed in 1984, and last updated in 1990. CDC and CYBER are trademarks of Control Data Corporation.

A Reynolds Number Study of Wing Leading-Edge Effects on a Supersonic Transport Model at Mach 0.3

NASA Technical Reports Server (NTRS)

Williams, M. Susan; Owens, Lewis R., Jr.; Chu, Julio

1999-01-01

A representative supersonic transport design was tested in the National Transonic Facility (NTF) in its original configuration with small-radius leading-edge flaps and also with modified large-radius inboard leading-edge flaps. Aerodynamic data were obtained over a range of Reynolds numbers at a Mach number of 0.3 and angles of attack up to 16 deg. Increasing the radius of the inboard leading-edge flap delayed nose-up pitching moment to a higher lift coefficient. Deflecting the large-radius leading-edge flap produced an overall decrease in lift coefficient and delayed nose-up pitching moment to even higher angles of attack as compared with the undeflected large- radius leading-edge flap. At angles of attack corresponding to the maximum untrimmed lift-to-drag ratio, lift and drag coefficients decreased while lift-to-drag ratio increased with increasing Reynolds number. At an angle of attack of 13.5 deg., the pitching-moment coefficient was nearly constant with increasing Reynolds number for both the small-radius leading-edge flap and the deflected large-radius leading-edge flap. However, the pitching moment coefficient increased with increasing Reynolds number for the undeflected large-radius leading-edge flap above a chord Reynolds number of about 35 x 10 (exp 6).

An aeroelastic instability provides a possible basis for the transition from gliding to flapping flight.

PubMed

Curet, Oscar M; Swartz, Sharon M; Breuer, Kenneth S

2013-03-06

The morphology, kinematics and stiffness properties of lifting surfaces play a key role in the aerodynamic performance of vertebrate flight. These surfaces, as a result of their flexible nature, may move both actively, owing to muscle contraction, and passively, in reaction to fluid forces. However, the nature and implications of this fluid-structure interaction are not well understood. Here, we study passive flight (flight with no active wing actuation) and explore a physical mechanism that leads to the emergence of a natural flapping motion. We model a vertebrate wing with a compliant shoulder and the ability to camber with an idealized physical model consisting of a cantilevered flat plate with a hinged trailing flap. We find that at low wind speed the wing is stationary, but at a critical speed the wing spontaneously flaps. The lift coefficient is significantly enhanced once the wing starts to oscillate, although this increase in lift generation is accompanied by an increase in drag. Flow visualization suggests that a strong leading edge vortex attached to the wing during downstroke is the primary mechanism responsible for the enhanced lift. The flapping instability we observe suggests a possible scenario for an evolutionary transition from gliding to powered flapping flight in animals that possess compliant wings capable of passive camber. Although the flapping state is accompanied by a lower lift-to-drag ratio, the increased lifting capability it confers might have enabled increased body mass, improved foraging performance and/or flight at lower speeds, any of which might have been selectively advantageous.

An aeroelastic instability provides a possible basis for the transition from gliding to flapping flight

PubMed Central

Curet, Oscar M.; Swartz, Sharon M.; Breuer, Kenneth S.

2013-01-01

The morphology, kinematics and stiffness properties of lifting surfaces play a key role in the aerodynamic performance of vertebrate flight. These surfaces, as a result of their flexible nature, may move both actively, owing to muscle contraction, and passively, in reaction to fluid forces. However, the nature and implications of this fluid–structure interaction are not well understood. Here, we study passive flight (flight with no active wing actuation) and explore a physical mechanism that leads to the emergence of a natural flapping motion. We model a vertebrate wing with a compliant shoulder and the ability to camber with an idealized physical model consisting of a cantilevered flat plate with a hinged trailing flap. We find that at low wind speed the wing is stationary, but at a critical speed the wing spontaneously flaps. The lift coefficient is significantly enhanced once the wing starts to oscillate, although this increase in lift generation is accompanied by an increase in drag. Flow visualization suggests that a strong leading edge vortex attached to the wing during downstroke is the primary mechanism responsible for the enhanced lift. The flapping instability we observe suggests a possible scenario for an evolutionary transition from gliding to powered flapping flight in animals that possess compliant wings capable of passive camber. Although the flapping state is accompanied by a lower lift-to-drag ratio, the increased lifting capability it confers might have enabled increased body mass, improved foraging performance and/or flight at lower speeds, any of which might have been selectively advantageous. PMID:23303221

Critical review of the trailing edge condition in steady and unsteady flow. Blade flutter in compressors and fans: Numerical simulation of the aerodynamic loading

NASA Technical Reports Server (NTRS)

Radwan, S. F.; Rockwell, D. O.; Johnson, S. H.

1982-01-01

Existing interpretations of the trailing edge condition, addressing both theoretical and experimental works in steady, as well as unsteady flows are critically reviewed. The work of Kutta and Joukowski on the trailing edge condition in steady flow is reviewed. It is shown that for most practical airfoils and blades (as in the case of most turbomachine blades), this condition is violated due to rounded trailing edges and high frequency effects, the flow dynamics in the trailing edge region being dominated by viscous forces; therefore, any meaningful modelling must include viscous effects. The question of to what extent the trailing edge condition affects acoustic radiation from the edge is raised; it is found that violation of the trailing edge condition leads to significant sound diffraction at the tailing edge, which is related to the problem of noise generation. Finally, various trailing edge conditions in unsteady flow are discussed, with emphasis on high reduced frequencies.

Initial Assessment of a Variable-Camber Continuous Trailing-Edge Flap System on a Rigid Wing for Drag Reduction in Subsonic Cruise

NASA Technical Reports Server (NTRS)

Ippolito, Corey; Nguyen, Nhan; Totah, Joe; Trinh, Khanh; Ting, Eric

2013-01-01

In this paper, we describe an initial optimization study of a Variable-Camber Continuous Trailing-Edge Flap (VCCTEF) system. The VCCTEF provides a light-weight control system for aircraft with long flexible wings, providing efficient high-lift capability for takeoff and landing, and greater efficiency with reduced drag at cruising flight by considering the effects of aeroelastic wing deformations in the control law. The VCCTEF system is comprised of a large number of distributed and individually-actuatable control surfaces that are constrained in movement relative to neighboring surfaces, and are non-trivially coupled through structural aeroelastic dynamics. Minimzation of drag results in a constrained, coupled, non-linear optimization over a high-dimension search space. In this paper, we describe the modeling, analysis, and optimization of the VCCTEF system control inputs for minimum drag in cruise. The purpose of this initial study is to quantify the expected benefits of the system concept. The scope of this analysis is limited to consideration of a rigid wing without structural flexibility in a steady-state cruise condition at various fuel weights. For analysis, we developed an optimization engine that couples geometric synthesis with vortex-lattice analysis to automate the optimization procedure. In this paper, we present and describe the VCCTEF system concept, optimization approach and tools, run-time performance, and results of the optimization at 20%, 50%, and 80% fuel load. This initial limited-scope study finds the VCCTEF system can potentially gain nearly 10% reduction in cruise drag, provides greater drag savings at lower operating weight, and efficiency is negatively impacted by the severity of relative constraints between control surfaces.

Application of Out-of-Plane Warping to Control Rotor Blade Twist

NASA Technical Reports Server (NTRS)

VanWeddingen, Yannick; Bauchau, Olivier; Kottapalli, Sesi; Ozbay, Serkan; Mehrotra, Yogesh

2012-01-01

The goal of this ongoing study is to develop and demonstrate the feasibility of a blade actuation system to dynamically change the twist, and/or the camber, of an airfoil section and, consequently, alter the in-flight aerodynamic loading on the blade for efficient flight control. The required analytical and finite element tools are under development to enable an accurate and comprehensive aeroelastic assessment of the current Full-Blade Warping and 3D Warping Actuated Trailing Edge Flap concepts. The feasibility of the current concepts for swashplateless rotors and higher harmonic blade control is also being investigated. In particular, the aim is to complete the following objectives, some of which have been completed (as noted below) and others that are currently ongoing: i) Develop a Vlasov finite element model and validate against the ABAQUS shell models (completed). ii) Implement the 3D warping actuation concept within the comprehensive analysis code DYMORE. iii) Perform preliminary aeroelastic simulations of blades using DYMORE with 3D warping actuation: a) Investigate the blade behavior under 1 per/rev actuation. Determine whether sufficient twist can be generated and sustained to achieve primary blade control. b) Investigate the behavior of a trailing edge flap configuration under higher harmonic excitations. Determine how much twist can be obtained at the harmonics 2-5 per/rev. iv) Determine actuator specifications such as the power required, load and displacements, and identify the stress and strain distributions in the actuated blades. In general, the completion of Item ii) above will give an additional research capability in rotorcraft dynamics analyses, i.e., the capability to calculate the rotor blade twist due to warping, something that is not currently available in any of the existing comprehensive rotorcraft analyses.

Wind Tunnel Results of the Aerodynamic Performance of a 1/8-Scale Model of a Twin-Engine Transport with Multi-Element Wing

NASA Technical Reports Server (NTRS)

Laflin, Brenda E. Gile; Applin, Zachary T.; Jones, Kenneth M.

1997-01-01

A wind tunnel investigation was performed in the 14- by 22-Foot Subsonic Tunnel on a pressure instrumented 1/8-scale twin-engine subsonic transport to better understand the flow physics on a multi-element wing section. The wing consisted of a part-span, triple-slotted trailing edge flap, inboard leading-edge Krueger flap and an outboard leading-edge slat. The model was instrumented with flush pressure ports at the fuselage centerline and seven spanwise wing locations. The model was tested in cruise, take-off and landing configurations at dynamic pressures and Mach numbers from 10 lbf/ft(exp 2) to 50 lbf/ft(exp 2) and 0.08 to 0.17, respectively. This resulted in corresponding Reynolds numbers of 0.8 x 10(exp 5) to 1.8 x 10(exp 6). Pressure data were collected using electronically scanned pressure devices and force and moment data were collected with a six component strain gauge balance. Results are presented for various control surface deflections over an angle-of-attack range from -4 degrees to 16 degrees and sideslip angle range from -10 degrees to 10 degrees. Longitudinal and lateral directional aerodynamic data are presented as well as chordwise pressure distributions at the seven spanwise wing locations and the fuselage centerline.

A numerical study of the effect of geometry variation, turbulence models, and dissipation on the flow past control surfaces

NASA Technical Reports Server (NTRS)

Chaussee, Denny S.

1993-01-01

The steady 3D viscous flow past the ONERA M6 wing and a slender delta wing-body with trailing edge control surfaces has been computed. A cell-centered finite-volume Navier-Stokes patched zonal method has been used for the numerical simulation. Both diagonalized and LUSGS schemes have been implemented. Besides the standard nonplanar zonal interfacing techniques, a new virtual zone capability has been employed. For code validation, the transonic flow past the ONERA M5 wing is calculated for angles-of-attack of 3.06 deg and 5.06 deg and compared with the available experiments. The wing-body computational results are compared with experimental data for both trailing-edge flaps deflected. The experimental flow conditions are M subinfinity = 0.4, a turbulent Reynolds number of 5.41 million based on a mean aerodynamic chord of 25.959 inches, adiabatic wall, and angles-of-attack varying from 0 deg to 23.85 deg. The computational results are presented for the 23.85 deg angle-of-attack case. The effects of the base flow due to a model sting, the varying second and fourth order numerical dissipation, and the turbulence model are all considered.

Structural and aerodynamic considerations of an active piezoelectric trailing-edge tab on a helicopter rotor

NASA Astrophysics Data System (ADS)

Murray, Gabriel Jon

This dissertation is concerned with an active tab for use on a rotorcraft for noise and vibration reduction. The tab is located at the trailing edge of the airfoil. The tab consists of a shim sandwiched by layers of the piezoelectric actuators, macro fiber composites, of varying length. This configuration is similar to a bimorph. The modus operandi is similar to that of a trailing edge flap. The actuators deform the tab, bending it to achieve a tip displacement. This provides a change in the lift, moment, and drag coefficients of the airfoil. By actuating the system at 3/rev to 5/rev, reductions in noise and vibration can be realized. The system was examined and designed around using the UH-60 Blackhawk as the model rotorcraft. The tab is envisioned to operate between 65% to 85% of the main rotor span. The tab's chordwise dimensions considered were 20% and 15% of the blade chord. In order to assess the potential of the tab to change the lift and moment coefficients of the airfoil-tab system, a steady computational fluid dynamics study was conducted. The results were generated via the University of Maryland's Transonic Unsteady Navier-Stokes code. Various tab deflection angles, Mach numbers, and angle-of-attack values were computed. These results were compared to a trailing edge flap of similar size. The comparison shows that the tab produces lift and moment increments similar to that of the trailing edge flap. The design of the tab---composed of both active piezoelectric actuators and passive materials---was conducted using finite element analysis. The objectives were to maximize the tip deflection due to the actuators, while minimizing the deformation due to inertial and aerodynamic forces and loads. The inertial loads (acceleration terms) come from both blade motion, such as flapping and pitch, as well as the rotation of the rotor (centrifugal force). All of these previously mentioned terms cause the tab to undergo undesirable deflections. The original concept consisted of a bimorph configuration with a single layer of macro fiber composite the entire length of the 20% chord tab. The final design has three, tapered actuator layers on either side of a shim. Also, the length of the tab was reduced to 15% of the chord. A multitude of designs were compared via their performance objectives as well as how the design variables changed performances relative to the objectives. The result of this detailed analysis was the selection of several configurations that were investigated in detail. With several designs selected, they formed the basis to build a prototype tab. This prototype was based on a 15% chord tab. Due to the available commercial actuator lengths, a two layer system was ultimately built. The tab tip displacement was measured with both static and harmonic inputs. Various input voltages, both with and without a DC offset, were tested, up to -500 V to 1500 V. The input frequency was varied between 0 Hz to 17.2 Hz. Also, a frequency response of the tab was generated. High voltage static tip displacement compares well with the expected result. The first bending natural frequency was measured at 138 Hz and compares well with the computed range of 135 Hz to 142 Hz, depending on bond layer thickness. The results of dynamic inputs uncovered an unexpected reduction in tip displacement of 60% to 65% of the static result, for high voltages. It is suspected that the actuator suffers from a time-dependent response, with quick initial displacement followed by a slow creeping to the final displacement value. This phenomena has also been observed in the literature, although it is not prevalent. Using the experiment as a guide to quantify the time-dependent behavior, the overall viability of the system is discussed. The maximum dynamic tip deflection (accounting for time-dependent behavior) of a 15% chord tab is 4°. The tab is able to generate lift coefficient increments up to 0.413 and moment increments of -0.068 for positive tab deflections. Deflections due to aerodynamic loads range from 0.3° to 2°. Deflections due to blade motions are less than 0.5° and due to the centrifugal force, 0.78°. This system can be stiffened, reducing the dynamic tip deflection to 3.25° and limiting the aerodynamic response to 1°. Overall, by designing the tab to modest deflections, moderate reductions in vibration are possible, but are limited by actuator authority. (Abstract shortened by UMI.)

A bio-inspired study on tidal energy extraction with flexible flapping wings.

PubMed

Liu, Wendi; Xiao, Qing; Cheng, Fai

2013-09-01

Previous research on the flexible structure of flapping wings has shown an improved propulsion performance in comparison to rigid wings. However, not much is known about this function in terms of power efficiency modification for flapping wing energy devices. In order to study the role of the flexible wing deformation in the hydrodynamics of flapping wing energy devices, we computationally model the two-dimensional flexible single and twin flapping wings in operation under the energy extraction conditions with a large Reynolds number of 106. The flexible motion for the present study is predetermined based on a priori structural result which is different from a passive flexibility solution. Four different models are investigated with additional potential local distortions near the leading and trailing edges. Our simulation results show that the flexible structure of a wing is beneficial to enhance power efficiency by increasing the peaks of lift force over a flapping cycle, and tuning the phase shift between force and velocity to a favourable trend. Moreover, the impact of wing flexibility on efficiency is more profound at a low nominal effective angle of attack (AoA). At a typical flapping frequency f * = 0.15 and nominal effective AoA of 10°, a flexible integrated wing generates 7.68% higher efficiency than a rigid wing. An even higher increase, around six times that of a rigid wing, is achievable if the nominal effective AoA is reduced to zero degrees at feathering condition. This is very attractive for a semi-actuated flapping energy system, where energy input is needed to activate the pitching motion. The results from our dual-wing study found that a parallel twin-wing device can produce more power compared to a single wing due to the strong flow interaction between the two wings.

Three-dimensional vortex wake structure of flapping wings in hovering flight.

PubMed

Cheng, Bo; Roll, Jesse; Liu, Yun; Troolin, Daniel R; Deng, Xinyan

2014-02-06

Flapping wings continuously create and send vortices into their wake, while imparting downward momentum into the surrounding fluid. However, experimental studies concerning the details of the three-dimensional vorticity distribution and evolution in the far wake are limited. In this study, the three-dimensional vortex wake structure in both the near and far field of a dynamically scaled flapping wing was investigated experimentally, using volumetric three-component velocimetry. A single wing, with shape and kinematics similar to those of a fruitfly, was examined. The overall result of the wing action is to create an integrated vortex structure consisting of a tip vortex (TV), trailing-edge shear layer (TESL) and leading-edge vortex. The TESL rolls up into a root vortex (RV) as it is shed from the wing, and together with the TV, contracts radially and stretches tangentially in the downstream wake. The downwash is distributed in an arc-shaped region enclosed by the stretched tangential vorticity of the TVs and the RVs. A closed vortex ring structure is not observed in the current study owing to the lack of well-established starting and stopping vortex structures that smoothly connect the TV and RV. An evaluation of the vorticity transport equation shows that both the TV and the RV undergo vortex stretching while convecting downwards: a three-dimensional phenomenon in rotating flows. It also confirms that convection and secondary tilting and stretching effects dominate the evolution of vorticity.

Study on Trailing Edge Ramp of Supercritical Airfoil

DTIC Science & Technology

2016-03-30

7 th Asia-Pacific International Symposium on Aerospace Technology, 25 – 27 November 2015, Cairns Study on Trailing Edge Ramp of Supercritical...China Abstract Trailing edge flow control method could improve the performance of supercritical airfoil with a small modification on the original...airfoil. In this paper, a ramp of 2%~7% chord length is sliced near the trailing edge to improve airfoil performance. The trailing edge ramp is

A Parametric Investigation of Nozzle Planform and Internal/External Geometry at Transonic Speeds

NASA Technical Reports Server (NTRS)

Cler, Daniel L.

1995-01-01

An experimental investigation of multidisciplinary (scarfed trailing edge) nozzle divergent flap geometry was conducted at transonic speeds in the NASA Langley 16-Foot Transonic Tunnel. The geometric parameters investigated include nozzle planform, nozzle contouring location (internal and/or external), and nozzle area ratio (area ratio 1.2 and 2.0). Data were acquired over a range of Mach Numbers from 0.6 to 1.2, angle-of-attack from 0.0 degrees to 9.6 degrees and nozzle pressure ratios from 1.0 to 20.0. Results showed that increasing the rate of change internal divergence angle across the width of the nozzle or increasing internal contouring will decrease static, aeropropulsive and thrust removed drag performance regardless of the speed regime. Also, increasing the rate of change in boattail angle across the width of the nozzle or increasing external contouring will provide the lowest thrust removed drag. Scarfing of the nozzle trailing edges reduces the aeropropulsive performance for the most part and adversely affects the nozzle plume shape at higher nozzle pressure ratios thus increasing the thrust removed drag. The effects of contouring were primary in nature and the effects of planform were secondary in nature. Larger losses occur supersonically than subsonically when scarfing of nozzle trailing edges occurs. The single sawtooth nozzle almost always provided lower thrust removed drag than the double sawtooth nozzles regardless the speed regime. If internal contouring is required, the double sawtooth nozzle planform provides better static and aeropropulsive performance than the single sawtooth nozzle and if no internal contouring is required the single sawtooth provides the highest static and aeropropulsive performance.

Survey of techniques for reduction of wind turbine blade trailing edge noise.

DOE Office of Scientific and Technical Information (OSTI.GOV)

Barone, Matthew Franklin

2011-08-01

Aerodynamic noise from wind turbine rotors leads to constraints in both rotor design and turbine siting. The primary source of aerodynamic noise on wind turbine rotors is the interaction of turbulent boundary layers on the blades with the blade trailing edges. This report surveys concepts that have been proposed for trailing edge noise reduction, with emphasis on concepts that have been tested at either sub-scale or full-scale. These concepts include trailing edge serrations, low-noise airfoil designs, trailing edge brushes, and porous trailing edges. The demonstrated noise reductions of these concepts are cited, along with their impacts on aerodynamic performance. Anmore » assessment is made of future research opportunities in trailing edge noise reduction for wind turbine rotors.« less

High-Fidelity Aerostructural Design Optimization of Transport Aircraft with Continuous Morphing Trailing Edge Technology

NASA Astrophysics Data System (ADS)

Burdette, David A., Jr.

Adaptive morphing trailing edge technology offers the potential to decrease the fuel burn of transonic commercial transport aircraft by allowing wings to dynamically adjust to changing flight conditions. Current configurations allow flap and aileron droop; however, this approach provides limited degrees of freedom and increased drag produced by gaps in the wing's surface. Leading members in the aeronautics community including NASA, AFRL, Boeing, and a number of academic institutions have extensively researched morphing technology for its potential to improve aircraft efficiency. With modern computational tools it is possible to accurately and efficiently model aircraft configurations in order to quantify the efficiency improvements offered by mor- phing technology. Coupled high-fidelity aerodynamic and structural solvers provide the capability to model and thoroughly understand the nuanced trade-offs involved in aircraft design. This capability is important for a detailed study of the capabilities of morphing trailing edge technology. Gradient-based multidisciplinary design opti- mization provides the ability to efficiently traverse design spaces and optimize the trade-offs associated with the design. This thesis presents a number of optimization studies comparing optimized config- urations with and without morphing trailing edge devices. The baseline configuration used throughout this work is the NASA Common Research Model. The first opti- mization comparison considers the optimal fuel burn predicted by the Breguet range equation at a single cruise point. This initial singlepoint optimization comparison demonstrated a limited fuel burn savings of less than 1%. Given the effectiveness of the passive aeroelastic tailoring in the optimized non-morphing wing, the singlepoint optimization offered limited potential for morphing technology to provide any bene- fit. To provide a more appropriate comparison, a number of multipoint optimizations were performed. With a 3-point stencil, the morphing wing burned 2.53% less fuel than its optimized non-morphing counterpart. Expanding further to a 7-point stencil, the morphing wing used 5.04% less fuel. Additional studies demonstrate that the size of the morphing device can be reduced without sizable performance reductions, and that as aircraft wings' aspect ratios increase, the effectiveness of morphing trailing edge devices increases. The final set of studies in this thesis consider mission analy- sis, including climb, multi-altitude cruise, and descent. These mission analyses were performed with a number of surrogate models, trained with O(100) optimizations. These optimizations demonstrated fuel burn reductions as large as 5% at off-design conditions. The fuel burn predicted by the mission analysis was up to 2.7% lower for the morphing wing compared to the conventional configuration.

PIV Measurements on a Blowing Flap

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Stead, Daniel J.

2004-01-01

PIV measurements of the flow in the region of a flap side edge are presented for several blowing flap configurations. The test model is a NACA 63(sub 2)-215 Hicks Mod-B main-element airfoil with a half-span Fowler flap. Air is blown from small slots located along the flap side edge on either the top, bottom or side surfaces. The test set up is described and flow measurements for a baseline and three blowing flap configurations are presented. The effects that the flap tip jets have on the structure of the flap side edge flow are discussed for each of the flap configurations tested. The results indicate that blowing air from a slot located along the top surface of the flap greatly weakened the top vortex system and pushed it further off the top surface. Blowing from the bottom flap surface kept the strong side vortex further outboard while blowing from the side surface only strengthened the vortex system or accelerated the merging of the side vortex to the flap top surface. It is concluded that blowing from the top or bottom surfaces of the flap may lead to a reduction of flap side edge noise.

An Investigation of Wing and Aileron Loads Due to Deflected Inboard and Outboard Ailerons on a 4-Percent-Thick 30 deg Sweptback Wing at Transonic Speeds

NASA Technical Reports Server (NTRS)

Whitcomb, Charles F.; Critzos, Chris C.; Brown, Philippa F.

1961-01-01

An investigation has been conducted in the Langley 16-foot transonic tunnel to determine the changes in wing loading characteristics due to deflections of a plain faired flap-type inboard aileron, a plain faired flap-type outboard aileron, and a slab-sided thickened trailing edge outboard aileron. The test wing was 4 percent thick and had 30 sweep of the quarter chord, an aspect ratio of 3.0, a taper ratio of 0.2, and NACA 65A004 airfoil sections. The loading characteristics of the deflected ailerons were also investigated. The model was a sting-mounted wing-body combination, and pressure measurements over one wing panel (exposed area) and the ailerons were obtained for angles of attack from 0 to 20 at deflections up to +/- 15 deg for Mach numbers between 0.80 and 1.03. The test Reynolds number based on the wing mean aerodynamic chord was about 7.4 x 10(exp 6). The results of the investigation indicated that positive deflection of the plain faired flap-type inboard aileron caused significant added loading over the wing sections outboard of the aileron at all Mach numbers for model angles of attack from 0 deg or 4 deg up to 12 deg. Positive deflection of the two outboard ailerons (plain faired and slab sided with thickened trailing edge) caused significant added loading over the wing sections inboard of the ailerons for different model angle-of-attack ranges at the several test Mach numbers. The loading shapes over the ailerons were irregular and would be difficult to predict from theoretical considerations in the transonic speed range. The longitudinal and lateral center-of-pressure locations for the ailerons varied only slightly with increasing angle of attack and/or Mach number. Generally, the negative slopes of the variations of aileron hinge-moment coefficient with aileron deflection for all three ailerons varied similarly with Mach number at the test angles of attack.

Development of an Integrated Nonlinear Aeroservoelastic Flight Dynamic Model of the NASA Generic Transport Model

NASA Technical Reports Server (NTRS)

Nguyen, Nhan; Ting, Eric

2018-01-01

This paper describes a recent development of an integrated fully coupled aeroservoelastic flight dynamic model of the NASA Generic Transport Model (GTM). The integrated model couples nonlinear flight dynamics to a nonlinear aeroelastic model of the GTM. The nonlinearity includes the coupling of the rigid-body aircraft states in the partial derivatives of the aeroelastic angle of attack. Aeroservoelastic modeling of the control surfaces which are modeled by the Variable Camber Continuous Trailing Edge Flap is also conducted. The R.T. Jones' method is implemented to approximate unsteady aerodynamics. Simulations of the GTM are conducted with simulated continuous and discrete gust loads..

Aircraft Flutter Testing

NASA Technical Reports Server (NTRS)

1997-01-01

Wilmer Reed gained international recognition for his innovative research, contributions and patented ideas relating to flutter and aeroelasticity of aerospace vehicles at Langley Research Center. In the early 1980's, Reed retired from Langley and joined the engineering staff of Dynamic Engineering Inc. While at DEI, Reed conceived and patented the DEI Flutter Exciter, now used world-wide in flight flutter testing of new or modified aircraft designs. When activated, the DEI Flutter Exciter alternately deflects the airstream upward and downward in a rapid manner, creating a force similar to that produced by an oscillating trailing edge flap. The DEI Flutter Exciter is readily adaptable to a variety of aircraft.

Impact of trailing edge shape on the wake and propulsive performance of pitching panels

NASA Astrophysics Data System (ADS)

Van Buren, T.; Floryan, D.; Brunner, D.; Senturk, U.; Smits, A. J.

2017-01-01

The effects of changing the trailing edge shape on the wake and propulsive performance of a pitching rigid panel are examined experimentally. The panel aspect ratio is AR=1 , and the trailing edges are symmetric chevron shapes with convex and concave orientations of varying degree. Concave trailing edges delay the natural vortex bending and compression of the wake, and the mean streamwise velocity field contains a single jet. Conversely, convex trailing edges promote wake compression and produce a quadfurcated wake with four jets. As the trailing edge shape changes from the most concave to the most convex, the thrust and efficiency increase significantly.

Reduction of Flap Side Edge Noise - the Blowing Flap

NASA Technical Reports Server (NTRS)

Hutcheson, Florence V.; Brooks, THomas F.

2005-01-01

A technique to reduce the noise radiating from a wing-flap side edge is being developed. As an airplane wing with an extended flap is exposed to a subsonic airflow, air is blown outward through thin rectangular chord-wise slots at various locations along the side edges and side surface of the flap to weaken and push away the vortices that originate in that region of the flap and are responsible for important noise emissions. Air is blown through the slots at up to twice the local flow velocity. The blowing is done using one or multiple slots, where a slot is located along the top, bottom or side surface of the flap along the side edge, or also along the intersection of the bottom (or top) and side surfaces.

Use Deflected Trailing Edge to Improve the Aerodynamic Performance and Develop Low Solidity LPT Cascade

NASA Astrophysics Data System (ADS)

Chao, Li; Peigang, Yan; Xiangfeng, Wang; Wanjin, Han; Qingchao, Wang

2017-08-01

This paper investigates the feasibility of improving the aerodynamic performance of low pressure turbine (LPT) blade cascades and developing low solidity LPT blade cascades through deflected trailing edge. A deflected trailing edge improved aerodynamic performance of both LPT blade cascades and low solidity LPT blade cascades. For standard solidity LPT cascades, deflecting the trailing edge can decrease the energy loss coefficient by 20.61 % for a Reynolds number (Re) of 25,000 and freestream turbulence intensities (FSTI) of 1 %. For a low solidity LPT cascade, aerodynamic performance was also improved by deflecting the trailing edge. Solidity of the LPT cascade can be reduced by 12.5 % for blades with a deflected trailing edge without a drop in efficiency. Here, the flow control mechanism surrounding a deflected trailing edge was also revealed.

Navier-Stokes Computations of a Wing-Flap Model With Blowing Normal to the Flap Surface

NASA Technical Reports Server (NTRS)

Boyd, D. Douglas, Jr.

2005-01-01

A computational study of a generic wing with a half span flap shows the mean flow effects of several blown flap configurations. The effort compares and contrasts the thin-layer, Reynolds averaged, Navier-Stokes solutions of a baseline wing-flap configuration with configurations that have blowing normal to the flap surface through small slits near the flap side edge. Vorticity contours reveal a dual vortex structure at the flap side edge for all cases. The dual vortex merges into a single vortex at approximately the mid-flap chord location. Upper surface blowing reduces the strength of the merged vortex and moves the vortex away from the upper edge. Lower surface blowing thickens the lower shear layer and weakens the merged vortex, but not as much as upper surface blowing. Side surface blowing forces the lower surface vortex farther outboard of the flap edge by effectively increasing the aerodynamic span of the flap. It is seen that there is no global aerodynamic penalty or benefit from the particular blowing configurations examined.

Flexible flapping wings with self-organized microwrinkles.

PubMed

Tanaka, Hiroto; Okada, Hiroyuki; Shimasue, Yosuke; Liu, Hao

2015-06-29

Bio-inspired flapping wings with a wrinkled wing membrane were designed and fabricated. The wings consist of carbon fibre-reinforced plastic frames and a polymer film with microscale wrinkles inspired by bird feathers and the corrugations of insect wings. The flexural and tensile stiffness of the wrinkled film can be controlled by modifying the orientations and waveforms of the wrinkles, thereby expanding the design space of flexible wings for micro flapping-wing aerial robots. A self-organization phenomenon was exploited in the fabrication of the microwrinkles such that microscale wrinkles spanning a broad wing area were spontaneously created. The wavy shape of these self-organized wrinkles was used as a mould, and a Parylene film was deposited onto the mould to form a wrinkled wing film. The effect of the waveforms of the wrinkles on the film stiffness was investigated theoretically, computationally and experimentally. Compared with a flat film, the flexural stiffness was increased by two orders of magnitude, and the tensile stiffness was reduced by two orders of magnitude. To demonstrate the effect of the wrinkles on the actual deformation of the flapping wings and the resulting aerodynamic forces, the fabricated wrinkled wings were tested using a tethered electric flapping mechanism. Chordwise unidirectional wrinkles were found to prevent fluttering near the trailing edge and to produce a greater aerodynamic lift compared with a flat wing or a wing with spanwise wrinkles. Our results suggest that the fine stiffness control of the wing film that can be achieved by tuning the microwrinkles can improve the aerodynamic performance of future flapping-wing aerial robots.

Aeroacoustic Measurements of a Wing-Flap Configuration

NASA Technical Reports Server (NTRS)

Meadows, Kristine R.; Brooks, Thomas F.; Humphreys, William M.; Hunter, William H.; Gerhold, Carl H.

1997-01-01

Aeroacoustic measurements are being conducted to investigate the mechanisms of sound generation in high-lift wing configurations, and initial results are presented. The model is approximately 6 percent of a full scale configuration, and consists of a main element NACA 63(sub 2) - 215 wing section and a 30 percent chord half-span flap. Flow speeds up to Mach 0.17 are tested at Reynolds number up to approximately 1.7 million. Results are presented for a main element at a 16 degree angle of attack, and flap deflection angles of 29 and 39 degrees. The measurement systems developed for this test include two directional arrays used to localize and characterize the noise sources, and an array of unsteady surface pressure transducers used to characterize wave number spectra and correlate with acoustic measurements. Sound source localization maps show that locally dominant noise sources exist on the flap-side edge. The spectral distribution of the noise sources along the flap-side edge shows a decrease in frequency of the locally dominant noise source with increasing distance downstream of the flap leading edge. Spectra are presented which show general spectral characteristics of Strouhal dependent flow-surface interaction noise. However, the appearance of multiple broadband tonal features at high frequency indicates the presence of aeroacoustic phenomenon following different scaling characteristics. The scaling of the high frequency aeroacoustic phenomenon is found to be different for the two flap deflection angles tested. Unsteady surface pressure measurements in the vicinity of the flap edge show high coherence levels between adjacent sensors on the flap-side edge and on the flap edge upper surface in a region which corresponds closely to where the flap-side edge vortex begins to spill over to the flap upper surface. The frequency ranges where these high levels of coherence occur on the flap surface are consistent with the frequency ranges in which dominant features appear in far field acoustic spectra. The consistency of strongly correlated unsteady surface pressures and far field pressure fluctuations suggests the importance of regions on the flap edge in generating sound.

Aeroacoustic analysis of an airfoil with Gurney flap based on time-resolved particle image velocimetry measurements

NASA Astrophysics Data System (ADS)

Zhang, Xueqing; Sciacchitano, Andrea; Pröbsting, Stefan

2018-05-01

Particle image velocimetry for the experimental assessment of trailing edge noise sources has become focus of research in recent years. The present study investigates the feasibility of the noise prediction for high-lift devices based on time-resolved particle image velocimetry (PIV). The model under investigation is a NACA 0015 airfoil with a Gurney flap with a height of 6% of the chord length. The velocity fields around and downstream of the Gurney flap were measured by PIV and used to compute the corresponding pressure fields by solving the Poisson equation for incompressible flows. The reconstructed pressure fluctuations on the airfoil surface constitute the source term for Curle's aeroacoustic analogy, which was employed in both the distributed and compact formulation to estimate the noise emission from PIV. The results of the two formulations are compared with the simultaneous far-field microphone measurements in the temporal and spectral domains. Both formulations of Curle's analogy yield acoustic sound pressure levels in good agreement with the simultaneous microphone measurements for the tonal component. The estimated far-field sound power spectra (SPL) from the PIV measurements reproduce the peak at the vortex shedding frequency, which also agrees well with the acoustic measurements.

Independent Orbiter Assessment (IOA): Analysis of the body flap subsystem

NASA Technical Reports Server (NTRS)

Wilson, R. E.; Riccio, J. R.

1986-01-01

The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items (PCIs). To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. The independent analysis results for the Orbiter Body Flap (BF) subsystem hardware are documented. The BF is a large aerosurface located at the trailing edge of the lower aft fuselage of the Orbiter. The proper function of the BF is essential during the dynamic flight phases of ascent and entry. During the ascent phase of flight, the BF trails in a fixed position. For entry, the BF provides elevon load relief, trim control, and acts as a heat shield for the main engines. Specifically, the BF hardware comprises the following components: Power Drive Unit (PDU), rotary actuators, and torque tubes. The IOA analysis process utilized available BF hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode. Of the 35 failure modes analyzed, 19 were determined to be PCIs.

Jet Flap Stator Blade Test in the High Reaction Turbine Blade Cascade Tunnel

NASA Image and Video Library

1970-03-21

A researcher examines the setup of a jet flap blade in the High Reaction Turbine Blade Cascade Tunnel at the National Aeronautics and Space Administration (NASA) Lewis Research Center. Lewis researchers were seeking ways to increase turbine blade loading on aircraft engines in an effort to reduce the overall size and weight of engines. The ability of each blade to handle higher loads meant that fewer stages and fewer blades were required. This study analyzed the performance of a turbine blade using a jet flap and high loading. A jet of air was injected into the main stream from the pressure surface near the trailing edge. The jet formed an aerodynamic flap which deflected the flow and changed the circulation around the blade and thus increased the blade loading. The air jet also reduced boundary layer thickness. The jet-flap blade design was appealing because the cooling air may also be used for the jet. The performance was studied in a two-dimensional cascade including six blades. The researcher is checking the jet flat cascade with an exit survey probe. The probe measured the differential pressure that was proportional to the flow angle. The blades were tested over a range of velocity ratios and three jet flow conditions. Increased jet flow improved the turning and decreased both the weight flow and the blade loading. However, high blade loadings were obtained at all jet flow conditions.

Effect of trailing edge shape on the wake and propulsive performance of pitching panels

NASA Astrophysics Data System (ADS)

van Buren, Tyler; Floryan, Daniel; Brunner, Daniel; Senturk, Utku; Smits, Alexander

2016-11-01

We present the effects of the trailing edge shape on the wake and propulsive performance of a pitching panel with an aspect ratio of 1. The trailing edges are symmetric chevron shapes with convex and concave orientations of varying degree. Concave trailing edges delay the natural vortex bending and compression of the wake, and the streamwise velocity field contains a single jet-like structure. Conversely, convex trailing edges promote wake compression and produce a wake split into four jets. Deviation from the square trailing edge mostly reduces the thrust and efficiency. Supported by the Office of Naval Research under MURI Grant Number N00014-14-1-0533.

Gas turbine bucket wall thickness control

DOEpatents

Stathopoulos, Dimitrios; Xu, Liming; Lewis, Doyle C.

2002-01-01

A core for use in casting a turbine bucket including serpentine cooling passages is divided into two pieces including a leading edge core section and a trailing edge core section. Wall thicknesses at the leading edge and the trailing edge of the turbine bucket can be controlled independent of each other by separately positioning the leading edge core section and the trailing edge core section in the casting die. The controlled leading and trailing edge thicknesses can thus be optimized for efficient cooling, resulting in more efficient turbine operation.

Sound suppression mixer

NASA Technical Reports Server (NTRS)

Brown, William H. (Inventor)

1994-01-01

A gas turbine engine flow mixer includes at least one chute having first and second spaced apart sidewalls joined together at a leading edge, with the sidewalls having first and second trailing edges defining therebetween a chute outlet. The first trailing edge is spaced longitudinally downstream from the second trailing edge for defining a septum in the first sidewall extending downstream from the second trailing edge. The septum includes a plurality of noise attenuating apertures.

Resonant Interaction of a Rectangular Jet with a Flat-Plate

NASA Technical Reports Server (NTRS)

Zaman, K. B. M. Q.; Fagan, A. F.; Clem, M. M.; Brown, C. A.

2014-01-01

A resonant interaction between a large aspect ratio rectangular jet and a flat-plate is addressed in this experimental study. The plate is placed parallel to but away from the direct path of the jet. At high subsonic conditions and for certain relative locations of the plate, the resonance accompanied by an audible tone is encountered. The trends of the tone frequency variation exhibit some similarities to, but also marked differences from, corresponding trends of the well-known edge-tone phenomenon. Under the resonant condition flow visualization indicates a periodic flapping motion of the jet column. Phase-averaged Mach number data obtained near the plate's trailing edge illustrate that the jet cross-section goes through large contortions within the period of the tone. Farther downstream a clear 'axis switching' takes place. These results suggest that the assumption of two-dimensionality should be viewed with caution in any analysis of the flow.

Buffet characteristics of the F-8 supercritical wing airplane

NASA Technical Reports Server (NTRS)

Deangelis, V. M.; Monaghan, R. C.

1977-01-01

The buffet characteristics of the F-8 supercritical wing airplane were investigated. Wing structural response was used to determine the buffet characteristics of the wing and these characteristics are compared with wind tunnel model data and the wing flow characteristics at transonic speeds. The wingtip accelerometer was used to determine the buffet onset boundary and to measure the buffet intensity characteristics of the airplane. The effects of moderate trailing edge flap deflections on the buffet onset boundary are presented. The supercritical wing flow characteristics were determined from wind tunnel and flight static pressure measurements and from a dynamic pressure sensor mounted on the flight test airplane in the vicinity of the shock wave that formed on the upper surface of the wing at transonic speeds. The comparison of the airplane's structural response data to the supercritical flow characteristics includes the effects of a leading edge vortex generator.

Effect of canard position and wing leading-edge flap deflection on wing buffet at transonic speeds

NASA Technical Reports Server (NTRS)

Gloss, B. B.; Henderson, W. P.; Huffman, J. K.

1974-01-01

A generalized wind-tunnel model, with canard and wing planform typical of highly maneuverable aircraft, was tested. The addition of a canard above the wing chord plane, for the configuration with leading-edge flaps undeflected, produced substantially higher total configuration lift coefficients before buffet onset than the configuration with the canard off and leading-edge flaps undeflected. The wing buffet intensity was substantially lower for the canard-wing configuration than the wing-alone configuration. The low-canard configuration generally displayed the poorest buffet characteristics. Deflecting the wing leading-edge flaps substantially improved the wing buffet characteristics for canard-off configurations. The addition of the high canard did not appear to substantially improve the wing buffet characteristics of the wing with leading-edge flaps deflected.

Flap Edge Aeroacoustic Measurements and Predictions

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.; Humphreys, William M., Jr.

2000-01-01

An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady surface pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that much of the physics is captured by the prediction models. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define the noise production. For the different edge conditions, extensive spectra and directivity are presented. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and surface roughness were successfully scaled by utilizing aerodynamic performance and boundary layer scaling methods developed herein.

Flap Edge Aeroacoustic Measurements and Predictions

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.; Humphreys, William M., Jr.

2000-01-01

An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a Small Aperture Directional Array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by Computational Fluid Dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady-surface-pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that much of the physics is captured by the prediction models. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define, the noise production. For the different edge conditions, extensive spectra and directivity are presented. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and surface roughness were successfully scaled by utilizing aerodynamic performance and boundary layer scaling method developed herein.

On the aerodynamic forces of flapping finite-wings in forward flight: a numerical study

NASA Astrophysics Data System (ADS)

Gonzalo, Alejandro; Uhlmann, Markus; Garcia-Villalba, Manuel; Flores, Oscar

2017-11-01

We study the flow around two flapping wings in forward flight at a low Reynolds number, Re = 500 , with 3D direct numerical simulations. The flow solver used is TUCAN, an in-house code which solves the Navier-Stokes equations for incompressible flow using an immersed boundary method to model the presence of the wings. The wings are rectangular with a NACA0012 airfoil of chord c as a cross-section. They are located side by side at a distance 0.5 c between their inboard tips. The wings flap with respect to an axis parallel to the streamwise velocity, without pitching. The angle of rotation is defined using a sinusoidal function with a reduced frequency k = 1 and an amplitude such that the maximum height of the outboard tips is c in all cases. We perform several simulations varying the aspect ratio of the wings (AR = 2 and 4) and the distance between the inboard tip of the wings and the axis of rotation (R = 0 , 2 and ∞), the latter case corresponding to wings in heaving motion. In this way we can study the variation of the fictitious forces on the wings and the induced spanwise flows, and their relation to the vortical structures on the wing (i.e. leading edge vortex, trailing edge votex, tip vortices) and the resulting aerodynamic forces. This work was funded by project TRA2013-41103-P (Mineco/Feder UE). The simulations were partially performed at the Steinbuch Centre for Computing, Karlsruhe, whose support is thankfully acknowledged.

Lift Optimization Study of a Multi-Element Three-Segment Variable Camber Airfoil

NASA Technical Reports Server (NTRS)

Kaul, Upender K.; Nguyen, Nhan T.

2016-01-01

This paper reports a detailed computational high-lift study of the Variable Camber Continuous Trailing Edge Flap (VCCTEF) system carried out to explore the best VCCTEF designs, in conjunction with a leading edge flap called the Variable Camber Krueger (VCK), for take-off and landing. For this purpose, a three-segment variable camber airfoil employed as a performance adaptive aeroelastic wing shaping control effector for a NASA Generic Transport Model (GTM) in landing and take-off configurations is considered. The objective of the study is to define optimal high-lift VCCTEF settings and VCK settings/configurations. A total of 224 combinations of VCK settings/configurations and VCCTEF settings are considered for the inboard GTM wing, where the VCCTEFs are configured as a Fowler flap that forms a slot between the VCCTEF and the main wing. For the VCK settings of deflection angles of 55deg, 60deg and 65deg, 18, 19 and 19 vck configurations, respectively, were considered for each of the 4 different VCCTEF deflection settings. Different vck configurations were defined by varying the horizontal and vertical distance of the vck from the main wing. A computational investigation using a Reynolds-Averaged Navier-Stokes (RANS) solver was carried out to complement a wind-tunnel experimental study covering three of these configurations with the goal of identifying the most optimal high-lift configurations. Four most optimal high-lift configurations, corresponding to each of the VCK deflection settings, have been identified out of all the different configurations considered in this study yielding the highest lift performance.

Computational Study of Porous Treatment for Altering Flap Side-Edge Flowfield

NASA Technical Reports Server (NTRS)

Choudhari, Meelan; Khorrami, Mehdi R.

2003-01-01

Reynolds-averaged Navier-Stokes calculations are used to investigate porous side-edge treatment as a passive means for flap noise reduction. Steady-state simulations are used to infer effects of the treatment on acoustically relevant features of the mean flow near the flap side edge. Application of the porous treatment over a miniscule fraction of the wetted flap area (scaling with the flap thickness) results in significantly weaker side-edge vortex structures via modification of the vortex initiation and roll-up processes. At high flap deflections, the region of axial flow reversal associated with the breakdown of the side-edge vortex is also eliminated, indicating an absence of vortex bursting in the presence of the treatment. Potential ramifications of the mean-flow modifications for flap-noise reduction are examined in the light of lessons learned from recent studies on flap noise. Computations confirm that any noise reduction benefit via the porous treatment would be achieved without compromising the aerodynamic effectiveness of the flap. Results of the parameter study contribute additional insight into the measured data from the 7x10 wind tunnel at NASA Ames and provide preliminary guidance for specifying optimal treatment characteristics in terms of treatment location, spatial extent, and flow resistance of the porous skin.

Partially turbulated trailing edge cooling passages for gas turbine nozzles

DOEpatents

Thatcher, Jonathan Carl; Burdgick, Steven Sebastian

2001-01-01

A plurality of passages are spaced one from the other along the length of a trailing edge of a nozzle vane in a gas turbine. The passages lie in communication with a cavity in the vane for flowing cooling air from the cavity through the passages through the tip of the trailing edge into the hot gas path. Each passage is partially turbulated and includes ribs in an aft portion thereof to provide enhanced cooling effects adjacent the tip of the trailing edge. The major portions of the passages are smooth bore. By this arrangement, reduced temperature gradients across the trailing edge metal are provided. Additionally, the inlets to each of the passages have a restriction whereby a reduced magnitude of compressor bleed discharge air is utilized for trailing edge cooling purposes.

KSC-2011-1201

NASA Image and Video Library

2011-01-25

CAPE CANAVERAL, Fla. -- Space shuttle Atlantis goes through a routine landing gear test in Orbiter Processing Facility-1 at NASA's Kennedy Space Center in Florida. Technicians are checking to make sure the shuttle's wheels, brakes, elevons and body flap function properly. Seen here, an elevon is tested. Elevons are located on the trailing edge of each wing and help control pitch of the shuttle as it comes in for landing. Atlantis is being prepared for the STS-135 mission, which will deliver the Raffaello multi-purpose logistics module packed with supplies, logistics and spare parts to the International Space Station. STS-135 is targeted to launch June 28, and will be the last spaceflight for the Space Shuttle Program. Photo credit: NASA/Jack Pfaller

KSC-2011-1202

NASA Image and Video Library

2011-01-25

CAPE CANAVERAL, Fla. -- Space shuttle Atlantis goes through a routine landing gear test in Orbiter Processing Facility-1 at NASA's Kennedy Space Center in Florida. Technicians are checking to make sure the shuttle's wheels, brakes, elevons and body flap function properly. Seen here, an elevon is tested. Elevons are located on the trailing edge of each wing and help control pitch of the shuttle as it comes in for landing. Atlantis is being prepared for the STS-135 mission, which will deliver the Raffaello multi-purpose logistics module packed with supplies, logistics and spare parts to the International Space Station. STS-135 is targeted to launch June 28, and will be the last spaceflight for the Space Shuttle Program. Photo credit: NASA/Jack Pfaller

Aircraft energy efficiency laminar flow control glove flight conceptual design study

NASA Technical Reports Server (NTRS)

Wright, A. S.

1979-01-01

A laminar flow control glove applied to the wing of a short to medium range jet transport with aft mounted engines was designed. A slotted aluminum glove concept and a woven stainless steel mesh porous glove concept suction surfaces were studied. The laminar flow control glove and a dummy glove with a modified supercritical airfoil, ducting, modified wing leading and trailing edges, modified flaps, and an LFC trim tab were applied to the wing after slot spacing suction parameters, and compression power were determined. The results show that a laminar flow control glove can be applied to the wing of a jet transport with an appropriate suction system installed.

Investigation of scrubbing and impingement noise

NASA Technical Reports Server (NTRS)

Fink, M. R.

1975-01-01

Tests were conducted in an acoustic wind tunnel to determine surface pressure spectra and far field noise caused by turbulence impinging on an airfoil and turbulence convected past a sharp trailing edge. Measured effects of flow velocity and turbulence intensity were compared with predictions from several theories. Also, tests were conducted in an anechoic chamber to determine surface pressure spectra and far field noise caused by a deflected airfoil scrubbed by a subsonic jet. This installation simulated both an under-the-wing and an upper-surface-blowing externally blown flap, depending on the deflection angle. Surface and far field spectra, and cross correlation coherence and delay time, were utilized to infer the major noise-producing mechanisms.

Flap-edge aeroacoustic measurements and predictions

NASA Astrophysics Data System (ADS)

Brooks, Thomas F.; Humphreys, William M.

2003-03-01

An aeroacoustic model test has been conducted to investigate the mechanisms of sound generation on high-lift wing configurations. This paper presents an analysis of flap side-edge noise, which is often the most dominant source. A model of a main element wing section with a half-span flap was tested at low speeds of up to a Mach number of 0.17, corresponding to a wing chord Reynolds number of approximately 1.7 million. Results are presented for flat (or blunt), flanged, and round flap-edge geometries, with and without boundary-layer tripping, deployed at both moderate and high flap angles. The acoustic database is obtained from a small aperture directional array (SADA) of microphones, which was constructed to electronically steer to different regions of the model and to obtain farfield noise spectra and directivity from these regions. The basic flap-edge aerodynamics is established by static surface pressure data, as well as by computational fluid dynamics (CFD) calculations and simplified edge flow analyses. Distributions of unsteady pressure sensors over the flap allow the noise source regions to be defined and quantified via cross-spectral diagnostics using the SADA output. It is found that shear layer instability and related pressure scatter is the primary noise mechanism. For the flat edge flap, two noise prediction methods based on unsteady-surface-pressure measurements are evaluated and compared to measured noise. One is a new causality spectral approach developed here. The other is a new application of an edge-noise scatter prediction method. The good comparisons for both approaches suggest that the prediction models capture much of the physics. Areas of disagreement appear to reveal when the assumed edge noise mechanism does not fully define the noise production. For the different edge conditions, extensive spectra and directivity are presented. The complexity of the directivity results demonstrate the strong role of edge source geometry and frequency in the noise radiation. Significantly, for each edge configuration, the spectra for different flow speeds, flap angles, and surface roughness were successfully scaled by utilizing aerodynamic performance and boundary-layer scaling methods developed herein.

Flutter suppression and gust alleviation using active controls

NASA Technical Reports Server (NTRS)

Nissim, E.

1975-01-01

Application of the aerodynamic energy approach to some problems of flutter suppression and gust alleviation were considered. A simple modification of the control-law is suggested for achieving the required pitch control in the use of a leading edge - trailing edge activated strip. The possible replacement of the leading edge - trailing edge activated strip by a trailing edge - tab strip is also considered as an alternate solution. Parameters affecting the performance of the activated leading edge - trailing edge strip were tested on the Arava STOL Transport and the Westwind Executive Jet Transport and include strip location, control-law gains and a variation in the control-law itself.

Trailing edge flow conditions as a factor in airfoil design

NASA Technical Reports Server (NTRS)

Ormsbee, A. I.; Maughmer, M. D.

1984-01-01

Some new developments relevant to the design of single-element airfoils using potential flow methods are presented. In particular, the role played by the non-dimensional trailing edge velocity in design is considered and the relationship between the specified value and the resulting airfoil geometry is explored. In addition, the ramifications of the unbounded trailing edge pressure gradients generally present in the potential flow solution of the flow over an airfoil are examined, and the conditions necessary to obtain a class of airfoils having finite trailing edge pressure gradients developed. The incorporation of these conditions into the inverse method of Eppler is presented and the modified scheme employed to generate a number of airfoils for consideration. The detailed viscous analysis of airfoils having finite trailing edge pressure gradients demonstrates a reduction in the strong inviscid-viscid interactions generally present near the trailing edge of an airfoil.

Experimental Study of Airfoil Trailing Edge Noise: Instrumentation, Methodology and Initial Results. M.S. Thesis

NASA Technical Reports Server (NTRS)

Manley, M. B.

1980-01-01

The mechanisms of aerodynamic noise generation at the trailing edge of an airfoil is investigated. Instrumentation was designed, a miniature semiconductor strain-gauge pressure transducer and associated electronic amplifier circuitry were designed and tested and digital signal analysis techniques applied to gain insight into the relationship between the dynamic pressure close to the trailing edge and the sound in the acoustic far-field. Attempts are made to verify some trailing-edge noise generation characteristics as theoretically predicted by several contemporary acousticians. It is found that the noise detected in the far-field is comprised of the sum of many uncorrelated emissions radiating from the vicinity of the trailing edge. These emissions appear to be the result of acoustic energy radiation which has been converted by the trailing-edge noise mechanism from the dynamic fluid energy of independent streamwise 'strips' of the turbulent boundary layer flow.

Flapping foil power generator performance enhanced with a spring-connected tail

NASA Astrophysics Data System (ADS)

Liu, Zhengliang; Tian, Fang-Bao; Young, John; Lai, Joseph C. S.

2017-12-01

The flexibility effects on the performance of a flapping foil power generator are numerically studied by using the immersed boundary-lattice Boltzmann method at a Reynolds number of 1100. The flapping foil system consists of a rigid NACA0015 foil undergoing harmonic pitch and plunge motions and a passively actuated flat plate pinned to the trailing edge of the rigid foil. The flexibility is modeled by a torsional spring model at the conjuncture of the rigid foil and the tail. Here, a parametric study on mass density and natural frequency is conducted under the optimum kinematic condition of the rigid system identified from the literature and numerical simulations made for reduced frequency f* = 0.04-0.24 and pitch amplitude θ0 = 40°-90°. Four typical cases are discussed in detail by considering time histories of hydrodynamic loads and tail deformations under the optimal and non-optimal kinematic conditions. Results show that under the rigid-system optimal kinematic condition, a tail with appropriate mass density (μ = 0.60) and resonant frequency ( fr*=1.18 ) can improve the maximum efficiency by 7.24% accompanied by an increase of 6.63% in power compared to those of a rigid foil with a rigid tail. This is because the deflection of the tail reduces the low pressure region on the pressure surface (i.e., the lower surface during the upstroke or the upper surface during the downstroke) caused by the leading edge vortex after the stroke reversal, resulting in a higher efficiency. At high flapping frequencies, a spring-connected tail ( fr*=0.13 ) eliminates the large spike in the moment observed in high stiffness cases, reducing the power required for the pitch motion, resulting in 117% improvement in efficiency over that with a rigid tail at a reduced frequency of 0.24.

Two-dimensional model of the interaction of a plane acoustic wave with nozzle edge and wing trailing edge.

PubMed

Faranosov, Georgy A; Bychkov, Oleg P

2017-01-01

The interaction of a plane acoustic wave with two-dimensional model of nozzle edge and trailing edge is investigated theoretically by means of the Wiener-Hopf technique. The nozzle edge and the trailing edge are simulated by two half-planes with offset edges. Shear layer behind the nozzle edge is represented by a vortex sheet supporting Kelvin-Helmholtz instability waves. The considered configuration combines two well-known models (nozzle edge and trailing edge), and reveals additional interesting physical aspects. To obtain the solution, the matrix Wiener-Hopf equation is solved in conjunction with a requirement that the full Kutta condition is imposed at the edges. Factorization of the kernel matrix is performed by the combination of Padé approximation and the pole removal technique. This procedure is used to obtain numerical results. The results indicate that the diffracted acoustic field may be significantly intensified due to scattering of hydrodynamic instability waves into sound waves provided that the trailing edge is close enough to the vortex sheet. Similar mechanism may be responsible for the intensification of jet noise near a wing.

Implementation of a Forth-Order Aeroelastic Coupling into a Viscous-Inviscid Flow Solver with Experimental Validation (for One Degree of Freedom)

NASA Astrophysics Data System (ADS)

Bartholomay, Sirko; Ramos-García, Néstor; Mikkelsen, Robert Flemming; Technical University of Denmark (DTU)-WInd Energy Team

2014-11-01

The viscous-inviscid flow solver Q3UIC for 2D aerodynamics has recently been developed at the Technical University of Denmark. The Q3UIC solver takes viscous and unsteady effects into account by coupling an unsteady inviscid panel method with the integral boundary layer equations by means of a strong coupling between the viscous and inviscid parts, and in this respect differs from other classic panel codes e.g. Xfoil. In the current work a Runge-Kutta-Nyström scheme was employed to couple inertial, elastic and aerodynamical forces and moments calculated by Q3UIC for a two-dimensional blade section in the time-domain. Numerical simulations are validated by a three step experimental verification process carried out in the low-turbulence wind tunnel at DTU. First, a comparison against steady experiments for a NACA 64418 profile and a flexible trailing edge flap is presented for different fixed flap angles, and second, the measured aerodynamic characteristics considering prescribed motion of the airfoil with a moving flap are compared to the Q3UIC predictions. Finally, an aeroelastic experiment for one degree of freedom-airfoil pitching- is used to evaluate the accuracy of aeroelastic coupling.

Experimental study on the use of synthetic jet actuators for lift control

NASA Astrophysics Data System (ADS)

Torres, Ricardo Benjamin

An experimental study on the use of synthetic jet actuators for lift control is conducted. The synthetic jet actuator is placed on the pressure side towards the trailing edge on a NACA 65(2)-415 airfoil representative of the cross section of an Inlet Guide Vane (IGV) in an industrial gas compressor. By redirecting or vectoring the shear layer at the trailing edge, the synthetic jet actuator increases lift and decreases drag on the airfoil without a mechanical device or flap. A compressor map that defines upper and lower bounds on operating velocities and airfoil dimensions, is compared with operating conditions of the low-speed wind tunnel at San Diego State University, to match gas compressor conditions in the wind tunnel. Realistic test conditions can range from Mach=0.12 to Mach= 0.27 and an airfoil chord from c=0.1 m to c=0.3 m. Based on the operating conditions, a final airfoil model is fabricated with a chord of c=0.1m. Several synthetic jet actuator designs are considered. A initial synthetic jet is designed to house a piezoelectric element with a material frequency of 1200 hz in a cavity with a volume of 4.47 cm3, a slot width of 0.25 mm, and a slot depth of 1.5 mm. With these dimensions, the Helmholtz frequency of the design is 1800Hz. Particle Image Velocimetry (PIV) experiments show that the design has a jet with a peak centerline jet velocity of 26 m/s at 750 Hz. A modified slant face synthetic jet is designed so that the cavity fits flush within the NACA airfoil surface. The slanted synthetic jet has a cavity volume of 4.67 cm3, a slot width of 0.25 mm, and a slot depth of 3.45 mm resulting in a Helmholtz frequency of 1170 hz for this design. PIV experiments show that the jet is redirected along the slant face according to the Coanda effect. A final synthetic jet actuator is directly integrated into the trailing edge of an airfoil with a cavity volume of 4.6 cm3, a slot width of 0.2 mm, and a slot depth of 1.6 mm. The Helmholtz frequency is 1450 Hz and matches closely with the piezoelectric element material frequency. The slot is designed so that actuator creates a jet normal to the airfoil surface. A wind tunnel model of the airfoil is 3D-printed with nine actuators integrated along the span of the airfoil. The synthetic jet slots cover 61% of the airfoil's span and the synthetic jet slots are located at a 13% chord upstream of the trailing edge. Tests are performed at multiple free stream velocities ranging from 17 m/s to 54 m/s which is the equivalent of an airfoil Reynolds number of Re=1.5105 to Re=4.5105. The integrated synthetic jet actuator increases lift. The increase is dependent on the freestream velocity, the actuation frequency, and angle of attack. For actuation at 1450 hz, and various freestream velocities, the synthetic jet actuator increases the lift by 2% at = alpha7° to 7% at = alpha15°. The synthetic jet increases L/D by 2% at = alpha7° to 15% at = alpha15°. Velocity contours obtained through PIV show that the synthetic jet turns the trailing edge shear layer similar to a Gurney flap, which increases lift. The synthetic jet reduces the wake velocity defect through injection of momentum, reducing the drag on the airfoil.

Clap-and-fling mechanism in a hovering insect-like two-winged flapping-wing micro air vehicle.

PubMed

Phan, Hoang Vu; Au, Thi Kim Loan; Park, Hoon Cheol

2016-12-01

This study used numerical and experimental approaches to investigate the role played by the clap-and-fling mechanism in enhancing force generation in hovering insect-like two-winged flapping-wing micro air vehicle (FW-MAV). The flapping mechanism was designed to symmetrically flap wings at a high flapping amplitude of approximately 192°. The clap-and-fling mechanisms were thereby implemented at both dorsal and ventral stroke reversals. A computational fluid dynamic (CFD) model was constructed based on three-dimensional wing kinematics to estimate the force generation, which was validated by the measured forces using a 6-axis load cell. The computed forces proved that the CFD model provided reasonable estimation with differences less than 8%, when compared with the measured forces. The measurement indicated that the clap and flings at both the stroke reversals augmented the average vertical force by 16.2% when compared with the force without the clap-and-fling effect. In the CFD simulation, the clap and flings enhanced the vertical force by 11.5% and horizontal drag force by 18.4%. The observations indicated that both the fling and the clap contributed to the augmented vertical force by 62.6% and 37.4%, respectively, and to the augmented horizontal drag force by 71.7% and 28.3%, respectively. The flow structures suggested that a strong downwash was expelled from the opening gap between the trailing edges during the fling as well as the clap at each stroke reversal. In addition to the fling phases, the influx of air into the low-pressure region between the wings from the leading edges also significantly contributed to augmentation of the vertical force. The study conducted for high Reynolds numbers also confirmed that the effect of the clap and fling was insignificant when the minimum distance between the two wings exceeded 1.2c (c = wing chord). Thus, the clap and flings were successfully implemented in the FW-MAV, and there was a significant improvement in the vertical force.

Clap-and-fling mechanism in a hovering insect-like two-winged flapping-wing micro air vehicle

PubMed Central

Phan, Hoang Vu; Au, Thi Kim Loan

2016-01-01

This study used numerical and experimental approaches to investigate the role played by the clap-and-fling mechanism in enhancing force generation in hovering insect-like two-winged flapping-wing micro air vehicle (FW-MAV). The flapping mechanism was designed to symmetrically flap wings at a high flapping amplitude of approximately 192°. The clap-and-fling mechanisms were thereby implemented at both dorsal and ventral stroke reversals. A computational fluid dynamic (CFD) model was constructed based on three-dimensional wing kinematics to estimate the force generation, which was validated by the measured forces using a 6-axis load cell. The computed forces proved that the CFD model provided reasonable estimation with differences less than 8%, when compared with the measured forces. The measurement indicated that the clap and flings at both the stroke reversals augmented the average vertical force by 16.2% when compared with the force without the clap-and-fling effect. In the CFD simulation, the clap and flings enhanced the vertical force by 11.5% and horizontal drag force by 18.4%. The observations indicated that both the fling and the clap contributed to the augmented vertical force by 62.6% and 37.4%, respectively, and to the augmented horizontal drag force by 71.7% and 28.3%, respectively. The flow structures suggested that a strong downwash was expelled from the opening gap between the trailing edges during the fling as well as the clap at each stroke reversal. In addition to the fling phases, the influx of air into the low-pressure region between the wings from the leading edges also significantly contributed to augmentation of the vertical force. The study conducted for high Reynolds numbers also confirmed that the effect of the clap and fling was insignificant when the minimum distance between the two wings exceeded 1.2c (c = wing chord). Thus, the clap and flings were successfully implemented in the FW-MAV, and there was a significant improvement in the vertical force. PMID:28083112

Assessment of a Conceptual Flap System Intended for Enhanced General Aviation Safety

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Carter, Melissa B.

2017-01-01

A novel multielement trailing-edge flap system for light general aviation airplanes was conceived for enhanced safety during normal and emergency landings. The system is designed to significantly reduce stall speed, and thus approach speed, with the goal of reducing maneuveringflight accidents and enhancing pilot survivability in the event of an accident. The research objectives were to assess the aerodynamic performance characteristics of the system and to evaluate the extent to which it provided both increased lift and increased drag required for the low-speed landing goal. The flap system was applied to a model of a light general aviation, high-wing trainer and tested in the Langley 12- Foot Low-Speed Wind Tunnel. Data were obtained for several device deflection angles, and component combinations at a dynamic pressure of 4 pounds per square foot. The force and moment data supports the achievement of the desired increase in lift with substantially increased drag, all at relatively shallow angles of attack. The levels of lift and drag can be varied through device deflection angles and inboard/outboard differential deflections. As such, it appears that this flap system may provide an enabling technology to allow steep, controllable glide slopes for safe rapid descent to landing with reduced stall speed. However, a simple flat-plate lower surface spoiler (LSS) provided either similar or superior lift with little impact on pitch or drag as compared to the proposed system. Higher-fidelity studies are suggested prior to use of the proposed system.

Active control system for a rotor blade trailing-edge flap

NASA Astrophysics Data System (ADS)

Duvernier, Marc; Reithler, Livier; Guerrero, Jean Y.; Rossi, Rinaldo A.

2000-06-01

Reducing the external noise is becoming a major issue for helicopter manufacturers. The idea beyond this goal is to reduce or even avoid the blade vortex interaction (BVI), especially during descent and flights over inhabited areas. This can be achieved by changing locally the lift of the blade. Several strategies to reach this goal are under investigation at EUROCOPTER such as the control of the local incidence of the blade by a direct lift flap. AEROSPATIALE MATRA Corporate Research Centre and AEROSPATIALE MATRA MISSILES proposed an actuator system able to answer EUROCOPTER's needs for moving a direct lift flap. The present paper describes the definition, manufacturing and testing of this new actuator system. This actuator is based on an electromagnetic patented actuation system developed by AEROSPATIALE MATRA MISSILES for missile and aeronautic applications. The particularity of this actuator is its ability to produce the desired force on its whole range of stroke. The flap is designed to be fitted on a DAUPHIN type blade produced by EUROCOPTER and the actuator system was designed to fit the room available within the blade and to produce the right amount of stroke and force within the required frequency range. Other constraints such as centrifugal loading were also taken into account. This paper describes briefly the specifications and the major characteristics of the actuating system and presents some results of its behavior on a representative composite test-bed manufactured by EUROCOPTER when subjected to realistic mechanical loads.

76 FR 27244 - Airworthiness Directives; The Boeing Company Model 747-400, 747-400D, and 747-400F Series...

Federal Register 2010, 2011, 2012, 2013, 2014

2011-05-11

... the flap control unit (FCU). This AD was prompted by a report of automatic retraction of the leading... takeoff. We are issuing this AD to prevent automatic retraction of the leading edge flaps during takeoff... automatic retraction of the leading edge flaps due to indications transmitted to the flap control unit (FCU...

Noise Testing of an Experimental Augmentor Wing

NASA Image and Video Library

1974-06-21

The augmentor wing concept was introduced during the early 1960s to enhance the performance of vertical and short takeoff (VSTOL) aircraft. The leading edge of the wing has full-span vertical flaps, and the trailing edge has double-slotted flaps. This provides aircraft with more control in takeoff and landing conditions. The augmentor wing also produced lower noise levels than other VSTOL designs. In the early 1970s Boeing Corporation built a Buffalo C-8A augmentor wing research aircraft for Ames Research Center. Researches at Lewis Research Center concentrated their efforts on reducing the noise levels of the wing. They initially used small-scale models to develop optimal nozzle screening methods. They then examined the nozzle designs on a large-scale model, seen here on an external test stand. This test stand included an airflow system, nozzle, the augmentor wing, and a muffler system below to reduce the atmospheric noise levels. The augmentor was lined with noise-reducing acoustic panels. The Lewis researchers were able to adjust the airflow to simulate conditions at takeoff and landing. Once the conditions were stabilized they took noise measurements from microphones placed in all directions from the wing, including an aircraft flying over. They found that the results coincided with the earlier small-scale studies for landing situations but not takeoffs. The acoustic panels were found to be successful.

Wing high-lift system with spoiler droop

NASA Astrophysics Data System (ADS)

Gubsky, Vitaly; Pavlenko, Olga; Petrov, Albert

2018-05-01

Take-off and landing regimes are becoming increasingly difficult due to the increased aircraft mass and speed and the toughening of air safety demands. The capabilities of conventional single-slotted or multi-slotted flaps are limited. This problem can be resolved by the deflection or bending of the tail section, which forms a potential flow that results in an increase in the lift component at high angles of attack. Different versions of the flap on a two-part high-lift device with a spoiler are examined. Calculations show that the application of an adaptive high-lift system with spoiler droop to an aircraft increases the lift coefficient by as much as 15%. Aircraft model experiments indicated that a δs = 5° deflection of the trailing edge increases the lift coefficient by 11% - 13%. Setting the propeller load coefficient to B = 1 (thrust/velocity factor propeller area) under take-off δf = 20° and landing δf = 35° configurations increases the efficiency of a lift coefficient by 25% and 36%, respectively. The application of spoiler droop and leading edge modifications of the wing expands the operational range of angles of attack and increases the lift coefficient by 10% at a small intensity (B = 0.5). It is concluded that the application of spoiler droop to high-lift system allows improvement of the wing properties under take-off and landing regimes of various types.

Peak-Seeking Optimization of Trim for Reduced Fuel Consumption: Architecture and Performance Predictions

NASA Technical Reports Server (NTRS)

Schaefer, Jacob; Brown, Nelson

2013-01-01

A peak-seeking control approach for real-time trim configuration optimization for reduced fuel consumption has been developed by researchers at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center to address the goals of the NASA Environmentally Responsible Aviation project to reduce fuel burn and emissions. The peak-seeking control approach is based on a steepest-descent algorithm using a time-varying Kalman filter to estimate the gradient of a performance function of fuel flow versus control surface positions. In real-time operation, deflections of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of an FA-18 airplane (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) are controlled for optimization of fuel flow. This presentation presents the design and integration of this peak-seeking controller on a modified NASA FA-18 airplane with research flight control computers. A research flight was performed to collect data to build a realistic model of the performance function and characterize measurement noise. This model was then implemented into a nonlinear six-degree-of-freedom FA-18 simulation along with the peak-seeking control algorithm. With the goal of eventual flight tests, the algorithm was first evaluated in the improved simulation environment. Results from the simulation predict good convergence on minimum fuel flow with a 2.5-percent reduction in fuel flow relative to the baseline trim of the aircraft.

Peak-Seeking Optimization of Trim for Reduced Fuel Consumption: Architecture and Performance Predictions

NASA Technical Reports Server (NTRS)

Schaefer, Jacob; Brown, Nelson A.

2013-01-01

A peak-seeking control approach for real-time trim configuration optimization for reduced fuel consumption has been developed by researchers at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center to address the goals of the NASA Environmentally Responsible Aviation project to reduce fuel burn and emissions. The peak-seeking control approach is based on a steepest-descent algorithm using a time-varying Kalman filter to estimate the gradient of a performance function of fuel flow versus control surface positions. In real-time operation, deflections of symmetric ailerons, trailing-edge flaps, and leading-edge flaps of an F/A-18 airplane (McDonnell Douglas, now The Boeing Company, Chicago, Illinois) are controlled for optimization of fuel flow. This paper presents the design and integration of this peak-seeking controller on a modified NASA F/A-18 airplane with research flight control computers. A research flight was performed to collect data to build a realistic model of the performance function and characterize measurement noise. This model was then implemented into a nonlinear six-degree-of-freedom F/A-18 simulation along with the peak-seeking control algorithm. With the goal of eventual flight tests, the algorithm was first evaluated in the improved simulation environment. Results from the simulation predict good convergence on minimum fuel flow with a 2.5-percent reduction in fuel flow relative to the baseline trim of the aircraft.

Equations and charts for the rapid estimation of hinge-moment and effectiveness parameters for trailing-edge controls having leading and trailing edges swept ahead of the Mach lines

NASA Technical Reports Server (NTRS)

Goin, Kennith L

1951-01-01

Existing conical-flow solutions have been used to calculate the hinge-moments and effectiveness parameters of trailing-edge controls having leading and trailing edges swept ahead of the Mach lines and having streamwise root and tip chords. Equations and detailed charts are presented for the rapid estimation of these parameters. Also included is an approximate method by which these parameters may be corrected for airfoil-section thickness.

Experimental Investigation of Aeroelastic Deformation of Slender Wings at Supersonic Speeds Using a Video Model Deformation Measurement Technique

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2013-01-01

A video-based photogrammetric model deformation system was established as a dedicated optical measurement technique at supersonic speeds in the NASA Langley Research Center Unitary Plan Wind Tunnel. This system was used to measure the wing twist due to aerodynamic loads of two supersonic commercial transport airplane models with identical outer mold lines but different aeroelastic properties. One model featured wings with deflectable leading- and trailing-edge flaps and internal channels to accommodate static pressure tube instrumentation. The wings of the second model were of single-piece construction without flaps or internal channels. The testing was performed at Mach numbers from 1.6 to 2.7, unit Reynolds numbers of 1.0 million to 5.0 million, and angles of attack from -4 degrees to +10 degrees. The video model deformation system quantified the wing aeroelastic response to changes in the Mach number, Reynolds number concurrent with dynamic pressure, and angle of attack and effectively captured the differences in the wing twist characteristics between the two test articles.

Ionic polymer-metal composite enabled robotic manta ray

NASA Astrophysics Data System (ADS)

Chen, Zheng; Um, Tae I.; Bart-Smith, Hilary

2011-04-01

The manta ray, Manta birostris, demonstrates excellent swimming capabilities; generating highly efficient thrust via flapping of dorsally flattened pectoral fins. In this paper, we present an underwater robot that mimics the swimming behavior of the manta ray. An assembly-based fabrication method is developed to create the artificial pectoral fins, which are capable of generating oscillatory with a large twisting angle between leading and trailing edges. Ionic polymer-metal composite (IPMC) actuators are used as artificial muscles in the fin. Each fin consists of four IPMC beams bonded with a compliant poly(dimethylsiloxane) (PDMS) membrane. By controlling each individual IPMC strips, we are able to generate complex flapping motions. The fin is characterized in terms of tip deflection, tip blocking force, twist angle, and power consumption. Based on the characteristics of the artificial pectoral fin, a small size and free-swimming robotic manta ray is developed. The robot consists of two artificial pectoral fins, a rigid body, and an on-board control unit with a lithium ion rechargeable battery. Experimental results show that the robot swam at a speed of up to 0.055 body length per second (BL/sec).

High-power piezo drive amplifier for large stack and PFC applications

NASA Astrophysics Data System (ADS)

Clingman, Dan J.; Gamble, Mike

2001-08-01

This paper describes the continuing development of Boeing High Power Piezo Drive Amplifiers. Described is the development and testing of a 1500 Vpp, 8 amp switching amplifier. This amplifier is used to drive a piezo stack driven rotor blade trailing edge flap on a full size helicopter. Also discuss is a switching amplifier designed to drive a Piezo Fiber Composite (PFC) active twist rotor blade. This amplifier was designed to drive the PFC material at 2000 Vpp and 0.5 amps. These amplifiers recycle reactive energy, allowing for a power and weight efficient amplifier design. This work was done in conjunction with the DARPA sponsored Phase II Smart Rotor Blade program and the NASA Langley Research Center sponsored Active Twist Rotor (ATR) blade program.

Recent advances in aerodynamic energy concept for flutter suppression and gust alleviation using active controls

NASA Technical Reports Server (NTRS)

Nissim, E.

1977-01-01

Control laws are derived, by using realizable transfer functions, which permit relaxation of the stability requirements of the aerodynamic energy concept. The resulting aerodynamic eigenvalues indicate that both the trailing edge and the leading edge-trailing edge control systems can be made more effective. These control laws permit the introduction of aerodynamic damping and stiffness terms in accordance with the requirements of any specific system. Flutter suppression and gust alleviation problems can now be treated by either a trailing edge control system or by a leading edge-trailing edge control system by using the aerodynamic energy concept. Results are applicable to a wide class of aircraft operating at subsonic Mach numbers.

Airfoil

DOEpatents

Ristau, Neil; Siden, Gunnar Leif

2015-07-21

An airfoil includes a leading edge, a trailing edge downstream from the leading edge, a pressure surface between the leading and trailing edges, and a suction surface between the leading and trailing edges and opposite the pressure surface. A first convex section on the suction surface decreases in curvature downstream from the leading edge, and a throat on the suction surface is downstream from the first convex section. A second convex section is on the suction surface downstream from the throat, and a first convex segment of the second convex section increases in curvature.

Numerical investigations on aerodynamic forces of deformable foils in hovering motions

NASA Astrophysics Data System (ADS)

Su, Xiaohui; Yin, Zhen; Cao, Yuanwei; Zhao, Yong

2017-04-01

In this paper, the aerodynamic forces of deformable foils for hovering flight are numerically investigated by a two-dimensional finite-volume arbitrary Lagrangian Eulerian Navier-Stokes solver. The effects of deformation on the lift force generation mechanisms of deformable wings in hovering flight are studied by comparison and analysis of deformable and rigid wing results. The prescribed deformation of the wings changes their morphing during hovering motion in both camber and angle of incidence. The effects of deflection amplitude, deflection phase, and rotation location on the aerodynamic performances of the foils, as well as the associated flow structures, are investigated in details, respectively. Results obtained show that foil morphing changes both Leading Edge Vortex (LEV) and Trailing Edge Vortex (TEV) generation and development processes. Consequently, the lift force generation mechanisms of deformable wings differ from those of rigid foil models. For the full deformation foil model studied, the effect of foil deformation enhances its lift force during both wake capture and delayed stall. There is an optimized camber amplitude, which was found to be 0.1*chord among those cases simulated. Partial deformation in the foil does not enhance its lift force due to unfavorable foil camber. TEV is significantly changed by the local angle of attack due to the foil deformation. On the other hand, Trailing Edge Flap (TEF) deflection in the hinge connected two-rigid-plate model directly affects the strength of both the LEV and TEV, thus influencing the entire vortex shedding process. It was found that lift enhancement can reach up to 33.5% just by the TEF deflection alone.

Flight test results from a supercritical mission adaptive wing with smooth variable camber

NASA Technical Reports Server (NTRS)

Powers, Sheryll Goecke; Webb, Lannie D.; Friend, Edward L.; Lokos, William A.

1992-01-01

The mission adaptive wing (MAW) consisted of leading- and trailing-edge variable-camber surfaces that could be deflected in flight to provide a near-ideal wing camber shape for any flight condition. These surfaces featured smooth, flexible upper surfaces and fully enclosed lower surfaces, distinguishing them from conventional flaps that have discontinuous surfaces and exposed or semiexposed mechanisms. Camber shape was controlled by either a manual or automatic flight control system. The wing and aircraft were extensively instrumented to evaluate the local flow characteristics and the total aircraft performance. This paper discusses the interrelationships between the wing pressure, buffet, boundary-layer and flight deflection measurement system analyses and describes the flight maneuvers used to obtain the data. The results are for a wing sweep of 26 deg, a Mach number of 0.85, leading and trailing-edge cambers (delta(sub LE/TE)) of 0/2 and 5/10, and angles of attack from 3.0 deg to 14.0 deg. For the well-behaved flow of the delta(sub LE/TE) = 0/2 camber, a typical cruise camber shape, the local and global data are in good agreement with respect to the flow properties of the wing. For the delta(sub LE/TE) = 5/10 camber, a maneuvering camber shape, the local and global data have similar trends and conclusions, but not the clear-cut agreement observed for cruise camber.

The effect of acoustic forcing on an airfoil tonal noise mechanism.

PubMed

Schumacher, Karn L; Doolan, Con J; Kelso, Richard M

2014-08-01

The response of the boundary layer over an airfoil with cavity to external acoustic forcing, across a sweep of frequencies, was measured. The boundary layer downstream of the cavity trailing edge was found to respond strongly and selectively at the natural airfoil tonal frequencies. This is considered to be due to enhanced feedback. However, the shear layer upstream of the cavity trailing edge did not respond at these frequencies. These findings confirm that an aeroacoustic feedback loop exists between the airfoil trailing edge and a location near the cavity trailing edge.

Flow-Field Investigation of Gear-Flap Interaction on a Gulfstream Aircraft Model

NASA Technical Reports Server (NTRS)

Yao, Chung-Sheng; Jenkins, Luther N.; Bartram, Scott M.; Harris, Jerome; Khorrami, Mehdi R.; Mace, W. Derry

2014-01-01

Off-surface flow measurements of a high-fidelity 18% scale Gulfstream aircraft model in landing configuration with the main landing gear deployed are presented. Particle Image Velocimetry (PIV) and Laser Velocimetry (LV) were used to measure instantaneous velocities in the immediate vicinity of the main landing gear and its wake and near the inboard tip of the flap. These measurements were made during the third entry of a series of tests conducted in the NASA Langley Research Center (LaRC) 14- by 22-Foot Subsonic Tunnel (14 x 22) to obtain a comprehensive set of aeroacoustic measurements consisting of both aerodynamic and acoustic data. The majority of the off-body measurements were obtained at a freestream Mach number of 0.2, angle of attack of 3 degrees, and flap deflection angle of 39 degrees with the landing gear on. A limited amount of data was acquired with the landing gear off. LV was used to measure the velocity field in two planes upstream of the landing gear and to measure two velocity profiles in the landing gear wake. Stereo and 2-D PIV were used to measure the velocity field over a region extending from upstream of the landing gear to downstream of the flap trailing edge. Using a special traverse system installed under the tunnel floor, the velocity field was measured at 92 locations to obtain a comprehensive picture of the pertinent flow features and characteristics. The results clearly show distinct structures in the wake that can be associated with specific components on the landing gear and give insight into how the wake is entrained by the vortex at the inboard tip of the flap.

Low-Order Modeling of Dynamic Stall on Airfoils in Incompressible Flow

NASA Astrophysics Data System (ADS)

Narsipur, Shreyas

Unsteady aerodynamics has been a topic of research since the late 1930's and has increased in popularity among researchers studying dynamic stall in helicopters, insect/bird flight, micro air vehicles, wind-turbine aerodynamics, and ow-energy harvesting devices. Several experimental and computational studies have helped researchers gain a good understanding of the unsteady ow phenomena, but have proved to be expensive and time-intensive for rapid design and analysis purposes. Since the early 1970's, the push to develop low-order models to solve unsteady ow problems has resulted in several semi-empirical models capable of effectively analyzing unsteady aerodynamics in a fraction of the time required by high-order methods. However, due to the various complexities associated with time-dependent flows, several empirical constants and curve fits derived from existing experimental and computational results are required by the semi-empirical models to be an effective analysis tool. The aim of the current work is to develop a low-order model capable of simulating incompressible dynamic-stall type ow problems with a focus on accurately modeling the unsteady ow physics with the aim of reducing empirical dependencies. The lumped-vortex-element (LVE) algorithm is used as the baseline unsteady inviscid model to which augmentations are applied to model unsteady viscous effects. The current research is divided into two phases. The first phase focused on augmentations aimed at modeling pure unsteady trailing-edge boundary-layer separation and stall without leading-edge vortex (LEV) formation. The second phase is targeted at including LEV shedding capabilities to the LVE algorithm and combining with the trailing-edge separation model from phase one to realize a holistic, optimized, and robust low-order dynamic stall model. In phase one, initial augmentations to theory were focused on modeling the effects of steady trailing-edge separation by implementing a non-linear decambering flap to model the effect of the separated boundary-layer. Unsteady RANS results for several pitch and plunge motions showed that the differences in aerodynamic loads between steady and unsteady flows can be attributed to the boundary-layer convection lag, which can be modeled by choosing an appropriate value of the time lag parameter, tau2. In order to provide appropriate viscous corrections to inviscid unsteady calculations, the non-linear decambering flap is applied with a time lag determined by the tau2 value, which was found to be independent of motion kinematics for a given airfoil and Reynolds number. The predictions of the aerodynamic loads, unsteady stall, hysteresis loops, and ow reattachment from the low-order model agree well with CFD and experimental results, both for individual cases and for trends between motions. The model was also found to perform as well as existing semi-empirical models while using only a single empirically defined parameter. Inclusion of LEV shedding capabilities and combining the resulting algorithm with phase one's trailing-edge separation model was the primary objective of phase two. Computational results at low and high Reynolds numbers were used to analyze the ow morphology of the LEV to identify the common surface signature associated with LEV initiation at both low and high Reynolds numbers and relate it to the critical leading-edge suction parameter (LESP ) to control the initiation and termination of LEV shedding in the low-order model. The critical LESP, like the tau2 parameter, was found to be independent of motion kinematics for a given airfoil and Reynolds number. Results from the final low-order model compared excellently with CFD and experimental solutions, both in terms of aerodynamic loads and vortex ow pattern predictions. Overall, the final combined dynamic stall model that resulted from the current research was successful in accurately modeling the physics of unsteady ow thereby helping restrict the number of empirical coefficients to just two variables while successfully modeling the aerodynamic forces and ow patterns in a simple and precise manner.

Numerical and experimental investigations on unsteady aerodynamics of flapping wings

NASA Astrophysics Data System (ADS)

Yu, Meilin

The development of a dynamic unstructured grid high-order accurate spectral difference (SD) method for the three dimensional compressible Navier-Stokes (N-S) equations and its applications in flapping-wing aerodynamics are carried out in this work. Grid deformation is achieved via an algebraic blending strategy to save computational cost. The Geometric Conservation Law (GCL) is imposed to ensure that grid deformation will not contaminate the flow physics. A low Mach number preconditioning procedure is conducted in the developed solver to handle the bio-inspired flow. The capability of the low Mach number preconditioned SD solver is demonstrated by a series of two dimensional (2D) and three dimensional (3D) simulations of the unsteady vortex dominated flow. Several topics in the flapping wing aerodynamics are numerically and experimentally investigated in this work. These topics cover some of the cutting-edge issues in flapping wing aerodynamics, including the wake structure analysis, airfoil thickness and kinematics effects on the aerodynamic performances, vortex structure analysis around 3D flapping wings and the kinematics optimization. Wake structures behind a sinusoidally pitching NACA0012 airfoil are studied with both experimental and numerical approaches. The experiments are carried out with Particle Image Velocimetry (PIV) and two types of wake transition processes, namely the transition from a drag-indicative wake to a thrust-indicative wake and that from the symmetric wake to the asymmetric wake are distinguished. The numerical results from the developed SD solver agree well with the experimental results. It is numerically found that the deflective direction of the asymmetric wake is determined by the initial conditions, e.g. initial phase angle. As most insects use thin wings (i. e., wing thickness is only a few percent of the chord length) in flapping flight, the effects of airfoil thickness on thrust generation are numerically investigated by simulating the flow fields around a series of plunging NACA symmetric airfoils with thickness ratio ranging from 4.0% to 20.0% of the airfoil chord length. The contribution of viscous force to flapping propulsion is accessed and it is found that viscous force becomes thrust producing, instead of drag producing, and plays a non-negligible role in thrust generation for thin airfoils. This is closely related to the variations of the dynamics of the unsteady vortex structures around the plunging airfoils. As nature flyers use complex wing kinematics in flapping flight, kinematics effects on the aerodynamic performance with different airfoil thicknesses are numerically studied by using a series of NACA symmetric airfoils. It is found that the combined plunging and pitching motion can outperform the pure plunging or pitching motion by sophisticatedly adjusting the airfoil gestures during the oscillation stroke. The thin airfoil better manipulates leading edge vortices (LEVs) than the thick airfoil (NACA0030) does in studied cases, and there exists an optimal thickness for large thrust generation with reasonable propulsive efficiency. With the present kinematics and dynamic parameters, relatively low reduced frequency is conducive for thrust production and propulsive efficiency for all tested airfoil thicknesses. In order to obtain the optimal kinematics parameters of flapping flight, a kinematics optimization is then performed. A gradient-based optimization algorithm is coupled with a second-order SD Navier-Stokes solver to search for the optimal kinematics of a certain airfoil undergoing a combined plunging and pitching motion. Then a high-order SD scheme is used to verify the optimization results and reveal the detailed vortex structures associated with the optimal kinematics of the flapping flight. It is found that for the case with maximum propulsive efficiency, there exists no leading edge separation during most of the oscillation cycle. In order to provide constructive suggestions to the design of micro-air-vehicles (MAVs), 3D simulations of the flapping wings are carried out in this work. Both the rectangular and bio-inspired wings with different kinematics are investigated. The formation process of two-jet-like wake patterns behind the finite-span flapping wing is found to be closely related to the interaction between trailing edge vortices and tip vortices. Then the effects of the wing planforms on the aerodynamics performance of the finite-span flapping wings are elucidated in terms of the evolution and dynamic interaction of unsteady vortex structures.

An experimental study of pressures on 60 deg Delta wings with leading edge vortex flaps

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Terry, J. E.; Donatelli, D. A.

1983-01-01

An experimental study was conducted in the Virginia Tech Stability Wind Tunnel to determine surface pressures over a 60 deg sweep delta wing with three vortex flap designs. Extensive pressure data was collected to provide a base data set for comparison with computational design codes and to allow a better understanding of the flow over vortex flaps. The results indicated that vortex flaps can be designed which will contain the leading edge vortex with no spillage onto the wing upper surface. However, the tests also showed that flaps designed without accounting for flap thickness will not be optimum and the result can be oversized flaps, early flap vortex reattachment and a second separation and vortex at the wing/flap hinge line.

Structural design of morphing trailing edge actuated by SMA

NASA Astrophysics Data System (ADS)

Wang, Qi; Xu, Zhiwei; Zhu, Qian

2013-09-01

In this paper, the morphing trailing edge is designed to achieve the up and down deflection under the aerodynamic load. After a detailed and accurate computational analysis to determine the SMA specifications and layout programs, a solid model is created in CATIA and the structures of the morphing wing trailing edge are produced by CNC machining. A set of DSP measurement and control system is designed to accomplish the controlling experiment of the morphing wing trailing edge. At last, via the force analysis, the trailing edge is fabricated with four sections of aluminum alloy, and the arrangement scheme of SMA wires is determined. Experiment of precise control integral has been performed to survey the control effect. The experiment consists of deflection angle tests of the third joint and the integral structure. Primarily, the ultimate deflection angle is tested in these two experiments. Therefore, the controlling experiment of different angles could be performed within this range. The results show that the deflection error is less than 4%and response time is less than 6.7 s, the precise controlling of the morphing trailing edge is preliminary realized.

Numerical analysis of the impact of permeability on trailing-edge noise

NASA Astrophysics Data System (ADS)

Koh, Seong Ryong; Meinke, Matthias; Schröder, Wolfgang

2018-05-01

The impact of porous surfaces on the near-wall turbulent structures and the generated trailing-edge noise is analyzed for several trailing-edge shapes of finite thickness using a high resolution large-eddy simulation (LES)/computational aeroacoustics (CAA) method. The porous surface of the trailing edge is defined by the porosity and the viscous permeability determined by the solution of a turbulent flat plate boundary layer at a Reynolds number 1280 based on the displacement thickness in the inflow cross section. The volume-averaged approach for the homogeneous porous medium shows that the porous impedance scales linearly with the porosity and exponentially with the mean structure size of a porous medium. The drag induced by the porous surface changes the friction velocity and the permeability Reynolds number ReK which determines the porous impedance Rs scaled by ReK-2/3. The trailing-edge noise is analyzed for three solid and three porous trailing edges. The effect of a finite span is investigated by the spanwise correlation model based on the measured coherence distribution. The acoustic prediction shows a good agreement with measurements of the broadband spectrum and the strong tone generated by a finite trailing-edge thickness. The pressure gradient inside the porous media is redistributed by the Darcy drag defined by the viscous permeability and the porosity. The mean pressure increases in the upstream direction inside the porous medium such that the flow acceleration involved in the acoustic generation is reduced inside the porous medium. The noise reduction by a porous medium reaches 11 dB for the trailing-edge shape which possesses a sharp corner for the solid surface. The porous surface applied to a semi-circular trailing edge achieves a 4 dB noise reduction. The directivity pattern for individual components of the acoustic spectrum shows that the massive noise reduction is determined at the tone. Enhanced wave diffraction by the thick flat plate changes the directivity pattern in the high frequency range.

Wind-tunnel studies of advanced cargo aircraft concepts. [leading edge vortex flaps for drag reduction

NASA Technical Reports Server (NTRS)

Rao, D. M.; Goglia, G. L.

1981-01-01

Accomplishments in vortex flap research are summarized. A singular feature of the vortex flap is that, throughout the range of angle of attack range, the flow type remains qualitatively unchanged. Accordingly, no large or sudden change in the aerodynamic characteristics, as happens when forcibly maintained attached flow suddenly reverts to separation, will occur with the vortex flap. Typical wind tunnel test data are presented which show the drag reduction potential of the vortex flap concept applied to a supersonic cruise airplane configuration. The new technology offers a means of aerodynamically augmenting roll-control effectiveness on slender wings at higher angles of attack by manipulating the vortex flow generated from leading edge separation. The proposed manipulator takes the form of a flap hinged at or close to the leading edge, normally retracted flush with the wing upper surface to conform to the airfoil shape.

Aerodynamics of dynamic wing flexion in translating wings

NASA Astrophysics Data System (ADS)

Liu, Yun; Cheng, Bo; Sane, Sanjay P.; Deng, Xinyan

2015-06-01

We conducted a systematic experimental study to investigate the aerodynamic effects of active trailing-edge flexion on a high-aspect-ratio wing translating from rest at a high angle of attack. We varied the timing and speed of the trailing-edge flexion and measured the resulting aerodynamic effects using a combination of direct force measurements and two-dimensional PIV flow measurements. The results indicated that the force and flow characteristics depend strongly on the timing of flexion, but relatively weakly on its speed. This is because the force and vortical flow structure are more sensitive to the timing of flexion relative to the shedding of starting vortex and leading-edge vortex. When the trailing-edge flexion occurred slightly before the starting vortex was shed, the lift production was greatly improved with the instantaneous peak lift increased by 54 % and averaged lift increased by 21 % compared with the pre-flexed case where the trailing-edge flexed before wing translation. However, when the trailing-edge flexed during or slightly after the leading-edge vortex shedding, the lift was significantly reduced by the disturbed development of leading-edge vortex. The force measurement results also imply that the trailing-edge flexion prior to wing translation does not augment lift but increases drag, thus resulting in a lower lift-drag ratio as compared to the case of flat wing.

Channel-wing System for Thrust Deflection and Force/Moment Generation

NASA Technical Reports Server (NTRS)

Englar, Robert J. (Inventor); Bushnell, Dennis M. (Inventor)

2006-01-01

An aircraft comprising a Channel Wing having blown c h - ne1 circulation control wings (CCW) for various functions. The blown channel CCW includes a channel that has a rounded or near-round trailing edge. The channel further has a trailing-edge slot that is adjacent to the rounded trailing edge of the channel. The trailing-edge slot has an inlet connected to a source of pressurized air and is capable of tangentially discharging pressurized air over the rounded trailing edge. The aircraft further has a propeller that is located in the channel and ahead of the rounded trailing edge of the channel. The propeller provides a propeller thrust exhaust stream across the channel wing to propel the aircraft through the air and to provide high lift. The pressurized air being discharged over the rounded trailing edge provides a high lift that is obtained independent of an aircraft angle of attack, thus preventing the asymmetry. separated flow, and stall experienced by the CC wing at the high angle of attack it required for high lift generation. The aircraft can further include blown outboard circulation control wings (CCW) that are synergistically connected to the blown channel CCWs. The blown outboard CCWs provide additional high lift, control thrust/drag interchange, and can provide all three aerodynamic moments when differential blowing is applied front-to-rear or left-to-right. Both the blown channel CCW and the outboard CCW also have leading-edge blowing slots to prevent flow separation or to provide aerodynamic moments for control.

Miniature Trailing Edge Effector for Aerodynamic Control

NASA Technical Reports Server (NTRS)

Lee, Hak-Tae (Inventor); Bieniawski, Stefan R. (Inventor); Kroo, Ilan M. (Inventor)

2008-01-01

Improved miniature trailing edge effectors for aerodynamic control are provided. Three types of devices having aerodynamic housings integrated to the trailing edge of an aerodynamic shape are presented, which vary in details of how the control surface can move. A bucket type device has a control surface which is the back part of a C-shaped member having two arms connected by the back section. The C-shaped section is attached to a housing at the ends of the arms, and is rotatable about an axis parallel to the wing trailing edge to provide up, down and neutral states. A flip-up type device has a control surface which rotates about an axis parallel to the wing trailing edge to provide up, down, neutral and brake states. A rotating type device has a control surface which rotates about an axis parallel to the chord line to provide up, down and neutral states.

Design and stable flight of a 21 g insect-like tailless flapping wing micro air vehicle with angular rates feedback control.

PubMed

Phan, Hoang Vu; Kang, Taesam; Park, Hoon Cheol

2017-04-04

An insect-like tailless flapping wing micro air vehicle (FW-MAV) without feedback control eventually becomes unstable after takeoff. Flying an insect-like tailless FW-MAV is more challenging than flying a bird-like tailed FW-MAV, due to the difference in control principles. This work introduces the design and controlled flight of an insect-like tailless FW-MAV, named KUBeetle. A combination of four-bar linkage and pulley-string mechanisms was used to develop a lightweight flapping mechanism that could achieve a high flapping amplitude of approximately 190°. Clap-and-flings at dorsal and ventral stroke reversals were implemented to enhance vertical force. In the absence of a control surface at the tail, adjustment of the location of the trailing edges at the wing roots to modulate the rotational angle of the wings was used to generate control moments for the attitude control. Measurements by a 6-axis load cell showed that the control mechanism produced reasonable pitch, roll and yaw moments according to the corresponding control inputs. The control mechanism was integrated with three sub-micro servos to realize the pitch, roll and yaw controls. A simple PD feedback controller was implemented for flight stability with an onboard microcontroller and a gyroscope that sensed the pitch, roll and yaw rates. Several flight tests demonstrated that the tailless KUBeetle could successfully perform a vertical climb, then hover and loiter within a 0.3 m ground radius with small variations in pitch and roll body angles.

A two-dimensional iterative panel method and boundary layer model for bio-inspired multi-body wings

NASA Astrophysics Data System (ADS)

Blower, Christopher J.; Dhruv, Akash; Wickenheiser, Adam M.

2014-03-01

The increased use of Unmanned Aerial Vehicles (UAVs) has created a continuous demand for improved flight capabilities and range of use. During the last decade, engineers have turned to bio-inspiration for new and innovative flow control methods for gust alleviation, maneuverability, and stability improvement using morphing aircraft wings. The bio-inspired wing design considered in this study mimics the flow manipulation techniques performed by birds to extend the operating envelope of UAVs through the installation of an array of feather-like panels across the airfoil's upper and lower surfaces while replacing the trailing edge flap. Each flap has the ability to deflect into both the airfoil and the inbound airflow using hinge points with a single degree-of-freedom, situated at 20%, 40%, 60% and 80% of the chord. The installation of the surface flaps offers configurations that enable advantageous maneuvers while alleviating gust disturbances. Due to the number of possible permutations available for the flap configurations, an iterative constant-strength doublet/source panel method has been developed with an integrated boundary layer model to calculate the pressure distribution and viscous drag over the wing's surface. As a result, the lift, drag and moment coefficients for each airfoil configuration can be calculated. The flight coefficients of this numerical method are validated using experimental data from a low speed suction wind tunnel operating at a Reynolds Number 300,000. This method enables the aerodynamic assessment of a morphing wing profile to be performed accurately and efficiently in comparison to Computational Fluid Dynamics methods and experiments as discussed herein.

Pressure investigation of NASA leading edge vortex flaps on a 60 deg Delta wing

NASA Technical Reports Server (NTRS)

Marchman, J. F., III; Donatelli, D. A.; Terry, J. E.

1983-01-01

Pressure distributions on a 60 deg Delta Wing with NASA designed leading edge vortex flaps (LEVF) were found in order to provide more pressure data for LEVF and to help verify NASA computer codes used in designing these flaps. These flaps were intended to be optimized designs based on these computer codes. However, the pressure distributions show that the flaps wre not optimum for the size and deflection specified. A second drag-producing vortex forming over the wing indicated that the flap was too large for the specified deflection. Also, it became apparent that flap thickness has a possible effect on the reattachment location of the vortex. Research is continuing to determine proper flap size and deflection relationships that provide well-behaved flowfields and acceptable hinge-moment characteristics.

7. Roaring Fork Motor Nature Trail, rocks along edge of ...

Library of Congress Historic Buildings Survey, Historic Engineering Record, Historic Landscapes Survey

7. Roaring Fork Motor Nature Trail, rocks along edge of road. - Great Smoky Mountains National Park Roads & Bridges, Roaring Fork Motor Nature Trail, Between Cherokee Orchard Road & U.S. Route 321, Gatlinburg, Sevier County, TN

On the aeroacoustic and flow structures developed on a flat plate with a serrated sawtooth trailing edge

NASA Astrophysics Data System (ADS)

Chong, Tze Pei; Vathylakis, Alexandros

2015-10-01

Results of an experimental study on turbulent flow over a flat plate with a serrated sawtooth trailing edge are presented in this paper. After tripping the boundary layer to become turbulent, the broadband noise sources at the sawtooth serrated trailing edge is studied by several experimental techniques. Broadband noise reduction by the serrated sawtooth trailing edge can be realistically achieved in the flat plate configuration. The variations of wall pressure power spectral density and the spanwise coherence (which relates to the spanwise correlation length) in a sawtooth trailing edge play a minor role in the mechanisms underpinning the reduction of self noise radiation. Conditional-averaging technique was applied in the boundary layer data where a pair of pressure-driven oblique vortical structures near the sawtooth side edges is identified. In the current flat plate configuration, the interaction between the vortical structures and the local turbulent boundary layer results in a redistribution of the momentum transport and turbulent shear stress near the sawtooth side edges as well as the sawtooth tip, thus affecting the efficiency of self noise radiation.

Flight test operations using an F-106B research airplane modified with a wing leading-edge vortex flap

NASA Technical Reports Server (NTRS)

Dicarlo, Daniel J.; Brown, Philip W.; Hallissy, James B.

1992-01-01

Flight tests of an F-106B aircraft equipped with a leading-edge vortex flap, which represented the culmination of a research effort to examine the effectiveness of the flap, were conducted at the NASA Langley Research Center. The purpose of the flight tests was to establish a data base on the use of a wing leading-edge vortex flap as a means to validate the design and analysis methods associated with the development of such a vortical flow-control concept. The overall experiment included: refinements of the design codes for vortex flaps; numerous wind tunnel entries to aid in verifying design codes and determining basic aerodynamic characteristics; design and fabrication of the flaps, structural modifications to the wing tip and leading edges of the test aircraft; development and installation of an aircraft research instrumentation system, including wing and flap surface pressure measurements and selected structural loads measurements; ground-based simulation to assess flying qualities; and finally, flight testing. This paper reviews the operational aspects associated with the flight experiment, which includes a description of modifications to the research airplane, the overall flight test procedures, and problems encountered. Selected research results are also presented to illustrate the accomplishments of the research effort.

Low-speed wind-tunnel tests of a one-tenth-scale model of a blended-arrow advanced supersonic transport. [conducted in Langley full-scale tunnel

NASA Technical Reports Server (NTRS)

Lemore, H. C.; Parett, L. P.

1975-01-01

Tests were conducted in the Langley full scale tunnel to determine the low-speed aerodynamic characteristics of a 1/10 scale model of a blended-arrow advanced supersonic transport. Tests were made for the clean configuration and a high-lift configuration with several combinations of leading- and trailing-edge flaps deflected for providing improved lift and longitudinal stability in the landing and takeoff modes. The tests were conducted for a range of angles of attack from about -6 deg to 30 deg, sideslip angles from -5 deg to 10 deg, and for Reynolds numbers from 6.78 x 1,000,000 to 13.85 x 1,000,000 corresponding to test velocities of 41 knots to 85 knots, respectively.

Real-Time Adaptive Least-Squares Drag Minimization for Performance Adaptive Aeroelastic Wing

NASA Technical Reports Server (NTRS)

Ferrier, Yvonne L.; Nguyen, Nhan T.; Ting, Eric

2016-01-01

This paper contains a simulation study of a real-time adaptive least-squares drag minimization algorithm for an aeroelastic model of a flexible wing aircraft. The aircraft model is based on the NASA Generic Transport Model (GTM). The wing structures incorporate a novel aerodynamic control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF). The drag minimization algorithm uses the Newton-Raphson method to find the optimal VCCTEF deflections for minimum drag in the context of an altitude-hold flight control mode at cruise conditions. The aerodynamic coefficient parameters used in this optimization method are identified in real-time using Recursive Least Squares (RLS). The results demonstrate the potential of the VCCTEF to improve aerodynamic efficiency for drag minimization for transport aircraft.

Integrated circuit cooled turbine blade

DOE Office of Scientific and Technical Information (OSTI.GOV)

Lee, Ching-Pang; Jiang, Nan; Um, Jae Y.

A turbine rotor blade includes at least two integrated cooling circuits that are formed within the blade that include a leading edge circuit having a first cavity and a second cavity and a trailing edge circuit that includes at least a third cavity located aft of the second cavity. The trailing edge circuit flows aft with at least two substantially 180-degree turns at the tip end and the root end of the blade providing at least a penultimate cavity and a last cavity. The last cavity is located along a trailing edge of the blade. A tip axial cooling channelmore » connects to the first cavity of the leading edge circuit and the penultimate cavity of the trailing edge circuit. At least one crossover hole connects the penultimate cavity to the last cavity substantially near the tip end of the blade.« less

Near wall cooling for a highly tapered turbine blade

DOEpatents

Liang, George [Palm City, FL

2011-03-08

A turbine blade having a pressure sidewall and a suction sidewall connected at chordally spaced leading and trailing edges to define a cooling cavity. Pressure and suction side inner walls extend radially within the cooling cavity and define pressure and suction side near wall chambers. A plurality of mid-chord channels extend radially from a radially intermediate location on the blade to a tip passage at the blade tip for connecting the pressure side and suction side near wall chambers in fluid communication with the tip passage. In addition, radially extending leading edge and trailing edge flow channels are located adjacent to the leading and trailing edges, respectively, and cooling fluid flows in a triple-pass serpentine path as it flows through the leading edge flow channel, the near wall chambers and the trailing edge flow channel.

Investigation of wave phenomena on a blunt airfoil with straight and serrated trailing edges

NASA Astrophysics Data System (ADS)

Nies, Juliane M.; Gageik, Manuel A.; Klioutchnikov, Igor; Olivier, Herbert

2015-07-01

An investigation of pressure waves in compressible subsonic and transonic flow around a generic airfoil is performed in a modified shock tube. New comprehensive results are presented on pressure waves in compressible flow. For the first time, the influence of trailing edge serration will be examined in terms of the reduction in pressure wave amplitude. A generic airfoil is tested in two main configurations, one with blunt trailing edges and the other one with serrated trailing edges in a Mach number range from 0.6 to 0.8 and at chord Reynolds numbers of 1 × 106 < Re c < 5 ×106. The flow of the blunt trailing edge is characterized by a regular vortex street in the wake creating a regular pattern of upstream-moving pressure waves along the airfoil. The observed pressure waves lead to strong pressure fluctuations within the local flow field. A reduction in the trailing edge thickness leads to a proportional increase in the frequency of the vortex street in the wake as well as the frequency of the waves deduced from constant Strouhal number. By serrating the trailing edge, the formation of vortices in the wake is disturbed. Therefore, also the upstream-moving waves are influenced and reduced in their strength resulting in a steadier flow. An increasing length of the saw tooth enhances the three dimensionality of the structures in the wake and causes a strong decrease in the wave amplitude.

Simulated big sagebrush regeneration supports predicted changes at the trailing and leading edges of distribution shifts

USGS Publications Warehouse

Schlaepfer, Daniel R.; Taylor, Kyle A.; Pennington, Victoria E.; Nelson, Kellen N.; Martin, Trace E.; Rottler, Caitlin M.; Lauenroth, William K.; Bradford, John B.

2015-01-01

Many semi-arid plant communities in western North America are dominated by big sagebrush. These ecosystems are being reduced in extent and quality due to economic development, invasive species, and climate change. These pervasive modifications have generated concern about the long-term viability of sagebrush habitat and sagebrush-obligate wildlife species (notably greater sage-grouse), highlighting the need for better understanding of the future big sagebrush distribution, particularly at the species' range margins. These leading and trailing edges of potential climate-driven sagebrush distribution shifts are likely to be areas most sensitive to climate change. We used a process-based regeneration model for big sagebrush, which simulates potential germination and seedling survival in response to climatic and edaphic conditions and tested expectations about current and future regeneration responses at trailing and leading edges that were previously identified using traditional species distribution models. Our results confirmed expectations of increased probability of regeneration at the leading edge and decreased probability of regeneration at the trailing edge below current levels. Our simulations indicated that soil water dynamics at the leading edge became more similar to the typical seasonal ecohydrological conditions observed within the current range of big sagebrush ecosystems. At the trailing edge, an increased winter and spring dryness represented a departure from conditions typically supportive of big sagebrush. Our results highlighted that minimum and maximum daily temperatures as well as soil water recharge and summer dry periods are important constraints for big sagebrush regeneration. Overall, our results confirmed previous predictions, i.e., we see consistent changes in areas identified as trailing and leading edges; however, we also identified potential local refugia within the trailing edge, mostly at sites at higher elevation. Decreasing regeneration probability at the trailing edge underscores the Schlaepfer et al. Future regeneration potential of big sagebrush potential futility of efforts to preserve and/or restore big sagebrush in these areas. Conversely, increasing regeneration probability at the leading edge suggest a growing potential for conflicts in management goals between maintaining existing grasslands by preventing sagebrush expansion versus accepting a shift in plant community composition to sagebrush dominance.

Noise measurements for various configurations of a model of a mixer nozzle externally blown flap system

NASA Technical Reports Server (NTRS)

Goodykoontz, J. H.; Wagner, J. M.; Sargent, N. B.

1973-01-01

Noise data were taken for variations to a large scale model of an externally blown flap lift augmentation system. The variations included two different mixer nozzles (7 and 8 lobes), two different wing models (2 and 3 flaps), and different lateral distances between the wing chord line and the nozzle centerline. When the seven lobe was used with the trailing flap in the 60 deg position, increasing the wing to nozzle distance had no effect on the sound level. When the eight lobe nozzle was used there was a decrease in sound level. With the 20 deg flap setting the noise level decreased when the distance was increased using either nozzle.

Wind Tunnel Investigation of Passive Porosity Applied to the Leading-Edge Extension and Leading-Edge Flaps on a Slender Wing at Subsonic Speed

NASA Technical Reports Server (NTRS)

Erickson, Gary E.

2017-01-01

A wind tunnel experiment was conducted in the NASA Langley Research Center 7- by 10-Foot High Speed Tunnel to determine the effects of passive surface porosity on the subsonic vortex flow interactions about a general research fighter configuration. Flow-through porosity was applied to the leading-edge extension, or LEX, and leading-edge flaps mounted to a 65deg cropped delta wing model as a potential vortex flow control technique at high angles of attack. All combinations of porous and nonporous LEX and flaps were investigated. Wing upper surface static pressure distributions and six-component forces and moments were obtained at a free-stream Mach number of 0.20 corresponding to a Reynolds number of 1.35(106) per foot, angles of attack up to 45deg, angles of sideslip of 0deg and +/-5deg, and leading-edge flap deflections of 0deg and 30deg.

Static Extended Trailing Edge for Lift Enhancement: Experimental and Computational Studies

DTIC Science & Technology

2007-06-01

3rd International Symposium on Integrating CFD and Experiments in Aerodynamics 20-21 June 2007 U.S. Air Force Academy, CO, USA Static Extended...is not significantly increased. Experiments and calculations are conducted to compare the aerodynamic characteristics of the extended trailing edge...basic configuration, has a good potential to improve the cruise flight efficiency. Key words: trailing edge, airfoil, wing, lift, drag, aerodynamics

Analytical observations on the aerodynamics of a delta wing with leading edge flaps

NASA Technical Reports Server (NTRS)

Oh, S.; Tavella, D.

1986-01-01

The effect of a leading edge flap on the aerodynamics of a low aspect ratio delta wing is studied analytically. The separated flow field about the wing is represented by a simple vortex model composed of a conical straight vortex sheet and a concentrated vortex. The analysis is carried out in the cross flow plane by mapping the wing trace, by means of the Schwarz-Christoffel transformation into the real axis of the transformed plane. Particular attention is given to the influence of the angle of attack and flap deflection angle on lift and drag forces. Both lift and drag decrease with flap deflection, while the lift-to-drag ratioe increases. A simple coordinate transformation is used to obtain a closed form expression for the lift-to-drag ratio as a function of flap deflection. The main effect of leading edge flap deflection is a partial suppression of the separated flow on the leeside of the wing. Qualitative comparison with experiments is presented, showing agreement in the general trends.

Wavy flow cooling concept for turbine airfoils

DOEpatents

Liang, George

2010-08-31

An airfoil including an outer wall and a cooling cavity formed therein. The cooling cavity includes a leading edge flow channel located adjacent a leading edge of the airfoil and a trailing edge flow channel located adjacent a trailing edge of the airfoil. Each of the leading edge and trailing edge flow channels define respective first and second flow axes located between pressure and suction sides of the airfoil. A plurality of rib members are located within each of the flow channels, spaced along the flow axes, and alternately extending from opposing sides of the flow channels to define undulating flow paths through the flow channels.

Comparing impacts between formal and informal recreational trails.

PubMed

Pickering, Catherine Marina; Norman, Patrick

2017-05-15

Globally there are hundreds of thousands of kilometres of recreational trails traversing natural areas of high conservation value: but what are their impacts and do impacts differ among trails? We compared the effects of four common types of recreational trails [(1) narrow and (2) medium width informal bare earth trails and (3) gravel and (4) tarmac/concrete formal trails] on vegetation adjacent to trails in a high conservation value plant community that is popular for mountain biking and hiking in Australia. Plant species composition was recorded in quadrats along the edge of the four types of trails and in control sites away from trails. Vegetation cover, the cover of individual growth forms, and species richness along the edges of all four types of trails were similar to the controls, although the wider trails affected plant composition, with the tarmac and gravel trails favouring different species. With very few comparative studies, more research is required to allow managers and researchers to directly compare differences in the severity and types of impacts on vegetation among trails. In the meantime, limiting damage to vegetation on the edge of hardened trails during construction, use and maintenance is important, and hardening trails may not always be appropriate. Copyright © 2016. Published by Elsevier Ltd.

Investigation of flow mechanism of a robotic fish swimming by using flow visualization synchronized with hydrodynamic force measurement

NASA Astrophysics Data System (ADS)

Tan, Guang-Kun; Shen, Gong-Xin; Huang, Shuo-Qiao; Su, Wen-Han; Ke, Yu

2007-11-01

When swimming in water by flapping its tail, a fish can overcome the drag from uniform flow and propel its body. The involved flow mechanism concerns 3-D and unsteady effects. This paper presents the investigation of the flow mechanism on the basis of a 3-D robotic fish model which has the typical geometry of body and tail with periodic flapping 2-freedom kinematical motion testing in the case of St = 0.78, Re = 6,600 and phase delay mode ( φ = -75°), in which may have a greater or maximum propulsion (without consideration of the optimal efficiency). Using a special technique of dye visualization which can clearly show vortex sheet and vortices in detail and using the inner 3-component force balance and cable supporting system with the phase-lock technique, the 3-D flow structure visualized in the wake of fish and the hydrodynamic force measurement were synchronized and obtained. Under the mentioned flapping parameters, we found the key flow structure and its evolution, a pair of complex 3-D chain-shape vortex (S-H vortex-rings, S1-H1 and S2-H2, and their legs L1 and L2) flow structures, which attach the leading edge and the trailing edge, then shed, move downstream and outwards and distribute two anti-symmetric staggering arrays along with the wake of the fish model in different phase stages during the flapping period. It is different with in the case of St = 0.25-0.35. Its typical flow structure and evolution are described and the results prove that they are different from the viewpoints based on the investigation of 2-D cases. For precision of the dynamic force measurement, in this paper it was provided with the method and techniques by subtracting the inertial forces and the forces induced by buoyancy and gravity effect in water, etc. from original data measured. The evolution of the synchronized measuring forces directly matching with the flow structure was also described in this paper.

Investigation of flow mechanism of a robotic fish swimming by using flow visualization synchronized with hydrodynamic force measurement

NASA Astrophysics Data System (ADS)

Tan, Guang-Kun; Shen, Gong-Xin; Huang, Shuo-Qiao; Su, Wen-Han; Ke, Yu

When swimming in water by flapping its tail, a fish can overcome the drag from uniform flow and propel its body. The involved flow mechanism concerns 3-D and unsteady effects. This paper presents the investigation of the flow mechanism on the basis of a 3-D robotic fish model which has the typical geometry of body and tail with periodic flapping 2-freedom kinematical motion testing in the case of St = 0.78, Re = 6,600 and phase delay mode (φ = - 75°), in which may have a greater or maximum propulsion (without consideration of the optimal efficiency). Using a special technique of dye visualization which can clearly show vortex sheet and vortices in detail and using the inner 3-component force balance and cable supporting system with the phase-lock technique, the 3-D flow structure visualized in the wake of fish and the hydrodynamic force measurement were synchronized and obtained. Under the mentioned flapping parameters, we found the key flow structure and its evolution, a pair of complex 3-D chain-shape vortex (S-H vortex-rings, S1 - H1 and S2 - H2, and their legs L1 and L2) flow structures, which attach the leading edge and the trailing edge, then shed, move downstream and outwards and distribute two antisymmetric staggering arrays along with the wake of the fish model in different phase stages during the flapping period. It is different with in the case of St = 0.25-0.35. Its typical flow structure and evolution are described and the results prove that they are different from the viewpoints based on the investigation of 2-D cases. For precision of the dynamic force measurement, in this paper it was provided with the method and techniques by subtracting the inertial forces and the forces induced by buoyancy and gravity effect in water, etc. from original data measured. The evolution of the synchronized measuring forces directly matching with the flow structure was also described in this paper.

Turbine blades and systems with forward blowing slots

DOEpatents

Zuteck, Michael D.; Zalusky, Leigh; Lees, Paul

2015-09-15

A blade for use in a wind turbine comprises a pressure side and suction side meeting at a trailing edge and leading edge. The pressure side and suction side provide lift to the turbine blade upon the flow of air from the leading edge to the trailing edge and over the pressure side and suction side. The blade includes one or more openings at the suction side, in some cases between the leading edge and the trailing edge. The one or more openings are configured to provide a pressurized fluid towards the leading edge of the blade, in some cases at an angle between about 0.degree. and 70.degree. with respect to an axis oriented from a centerline of the blade toward the leading edge.

Rotorcraft In-Plane Noise Reduction Using Active/Passive Approaches with Induced Vibration Tracking

NASA Astrophysics Data System (ADS)

Chia, Miang Hwee

A comprehensive study of the use of active and passive approaches for in-plane noise reduction, including the vibrations induced during noise reduction, was conducted on a hingeless rotor configuration resembling the MBB BO-105 rotor. First, a parametric study was performed to examine the effects of rotor blade stiffness on the vibration and noise reduction performance of a 20%c plain trailing edge flap and a 1.5%c sliding microflap. This was accomplished using a comprehensive code AVINOR (for Active VIbration and NOise Reduction). A two-dimensional unsteady reduced order aerodynamic model (ROM), using the Rational Function Approximation approach and CFD-based oscillatory aerodynamic load data, was used in the comprehensive code. The study identified a hingeless blade configuration with torsional frequency of 3.17/rev as an optimum configuration for studying vibration and noise reduction using on-blade control devices such as flaps or microflaps. Subsequently, a new suite of computational tools capable of predicting in-plane low frequency sound pressure level (LFSPL) rotorcraft noise and its control was developed, replacing the acoustic module WOPWOP in AVINOR with a new acoustic module HELINOIR (for HELIcopter NOIse Reduction), which overcomes certain limitations associated with WOPWOP. The new suite, consisting of the AVINOR/HELINOIR combination, was used to study active flaps, as well as microflaps operating in closed-loop mode for in-plane noise reduction. An alternative passive in-plane noise reduction approach using modification to the blade tip in the 10%R outboard region was also studied. The new suite consisting of the AVINOR/HELINOIR combination based on a compact aeroacoustic model was validated by comparing with wind tunnel test results, and subsequently verified by comparing with computational results. For active control, the in-plane noise reduction obtained with a single 20%c plain trailing edge flap during level flight at a moderate advance ratio was examined. Different configurations of far-field and near-field feedback microphone locations were examined to develop a fundamental understanding of the feedback microphone locations on the noise reduction process A near-field microphone located on the tip of a nose boom was found to produce a LFSPL reduction of up to 6dB. However, this noise reduction was accompanied by an out-of-plane noise increase of 18dB and 60% increase in vertical hub shear. For passive control, three tip geometries having sweep, dihedral, and anhedral, were considered. The tip dihedral reduced LFSPL by up to 2dB without a vibratory load penalty. However, this was accompanied by an increase in the mid frequency sound pressure levels (MFSPL). The tip sweep and tip anhedral produced an increase in in-plane LFSPL below the horizon. A comparison of the active and passive approaches indicated that active approaches implemented by a plain flap with a feedback microphone located on the nose boom is superior to the passive control approaches. However, there is a general trade-off between LFSPL reduction, MFSPL generation and vibratory hub loads induced by noise control.

Deformable wing kinematics in the desert locust: how and why do camber, twist and topography vary through the stroke?

PubMed Central

Walker, Simon M.; Thomas, Adrian L. R.; Taylor, Graham K.

2009-01-01

Here, we present a detailed analysis of the wing kinematics and wing deformations of desert locusts (Schistocerca gregaria, Forskål) flying tethered in a wind tunnel. We filmed them using four high-speed digital video cameras, and used photogrammetry to reconstruct the motion of more than 100 identified points. Whereas the hindwing motions were highly stereotyped, the forewing motions showed considerable variation, consistent with a role in flight control. Both wings were positively cambered on the downstroke. The hindwing was cambered through an ‘umbrella effect’ whereby the trailing edge tension compressed the radial veins during the downstroke. Hindwing camber was reversed on the upstroke as the wing fan corrugated, reducing the projected area by 30 per cent, and releasing the tension in the trailing edge. Both the wings were strongly twisted from the root to the tip. The linear decrease in incidence along the hindwing on the downstroke precisely counteracts the linear increase in the angle of attack that would otherwise occur in root flapping for an untwisted wing. The consequent near-constant angle of attack is reminiscent of the optimum for a propeller of constant aerofoil section, wherein a linear twist distribution allows each section to operate at the unique angle of attack maximizing the lift to drag ratio. This implies tuning of the structural, morphological and kinematic parameters of the hindwing for efficient aerodynamic force production. PMID:19091683

Numerical Investigations of the Influence of Unsteady Vane Trailing Edge Shock Wave on Film Cooling Effectiveness of Rotor Blade Leading Edge

NASA Astrophysics Data System (ADS)

Wang, Yufeng; Cai, Le; Wang, Songtao; Zhou, Xun

2018-04-01

Unsteady numerical simulations of a high-load transonic turbine stage have been carried out to study the influences of vane trailing edge outer-extending shockwave on rotor blade leading edge film cooling performance. The turbine stage used in this paper is composed of a vane section and a rotor one which are both near the root section of a transonic high-load turbine stage. The Mach number is 0.94 at vane outlet, and the relative Mach number is above 1.10 at rotor outlet. Various positions and oblique angles of film cooling holes were investigated in this research. Results show that the cooling efficiency on the blade surface of rotor near leading edge is significantly affected by vane trailing edge outer-extending shockwave in some cases. In the cases that film holes are close to leading edge, cooling performance suffers more from the sweeping vane trailing edge outer-extending shockwave. In addition, coolant flow ejected from oblique film holes is harder to separate from the blade surface of rotor, and can cover more blade area even under the effects of sweeping vane trailing edge shockwave. As a result, oblique film holes can provide better film cooling performance than vertical film holes do near the leading edge on turbine blade which is swept by shockwaves.

Validation and Analysis of Numerical Results for a Two-Pass Trapezoidal Channel With Different Cooling Configurations of Trailing Edge.

PubMed

Siddique, Waseem; El-Gabry, Lamyaa; Shevchuk, Igor V; Fransson, Torsten H

2013-01-01

High inlet temperatures in a gas turbine lead to an increase in the thermal efficiency of the gas turbine. This results in the requirement of cooling of gas turbine blades/vanes. Internal cooling of the gas turbine blade/vanes with the help of two-pass channels is one of the effective methods to reduce the metal temperatures. In particular, the trailing edge of a turbine vane is a critical area, where effective cooling is required. The trailing edge can be modeled as a trapezoidal channel. This paper describes the numerical validation of the heat transfer and pressure drop in a trapezoidal channel with and without orthogonal ribs at the bottom surface. A new concept of ribbed trailing edge has been introduced in this paper which presents a numerical study of several trailing edge cooling configurations based on the placement of ribs at different walls. The baseline geometries are two-pass trapezoidal channels with and without orthogonal ribs at the bottom surface of the channel. Ribs induce secondary flow which results in enhancement of heat transfer; therefore, for enhancement of heat transfer at the trailing edge, ribs are placed at the trailing edge surface in three different configurations: first without ribs at the bottom surface, then ribs at the trailing edge surface in-line with the ribs at the bottom surface, and finally staggered ribs. Heat transfer and pressure drop is calculated at Reynolds number equal to 9400 for all configurations. Different turbulent models are used for the validation of the numerical results. For the smooth channel low-Re k-ɛ model, realizable k-ɛ model, the RNG k-ω model, low-Re k-ω model, and SST k-ω models are compared, whereas for ribbed channel, low-Re k-ɛ model and SST k-ω models are compared. The results show that the low-Re k-ɛ model, which predicts the heat transfer in outlet pass of the smooth channels with difference of +7%, underpredicts the heat transfer by -17% in case of ribbed channel compared to experimental data. Using the same turbulence model shows that the height of ribs used in the study is not suitable for inducing secondary flow. Also, the orthogonal rib does not strengthen the secondary flow rotational momentum. The comparison between the new designs for trailing edge shows that if pressure drop is acceptable, staggered arrangement is suitable for the outlet pass heat transfer. For the trailing edge wall, the thermal performance for the ribbed trailing edge only was found about 8% better than other configurations.

Owl-inspired leading-edge serrations play a crucial role in aerodynamic force production and sound suppression.

PubMed

Rao, Chen; Ikeda, Teruaki; Nakata, Toshiyuki; Liu, Hao

2017-07-04

Owls are widely known for silent flight, achieving remarkably low noise gliding and flapping flights owing to their unique wing morphologies, which are normally characterized by leading-edge serrations, trailing-edge fringes and velvet-like surfaces. How these morphological features affect aerodynamic force production and sound suppression or noise reduction, however, is still not well known. Here we address an integrated study of owl-inspired single feather wing models with and without leading-edge serrations by combining large-eddy simulations (LES) with particle-image velocimetry (PIV) and force measurements in a low-speed wind tunnel. With velocity and pressure spectra analysis, we demonstrate that leading-edge serrations can passively control the laminar-turbulent transition over the upper wing surface, i.e. the suction surface at all angles of attack (0°  <  AoA  <  20°), and hence play a crucial role in aerodynamic force and sound production. We find that there exists a tradeoff between force production and sound suppression: serrated leading-edges reduce aerodynamic performance at lower AoAs  <  15° compared to clean leading-edges but are capable of achieving both noise reduction and aerodynamic performance at higher AoAs  >  15° where owl wings often reach in flight. Our results indicate that the owl-inspired leading-edge serrations may be a useful device for aero-acoustic control in biomimetic rotor designs for wind turbines, aircrafts, multi-rotor drones as well as other fluid machinery.

Helicopter rotor trailing edge noise. [noise prediction

NASA Technical Reports Server (NTRS)

Schlinker, R. H.; Amier, R. K.

1981-01-01

A two dimensional section of a helicopter main rotor blade was tested in an acoustic wind tunnel at close to full-scale Reynolds numbers to obtain boundary layer data and acoustic data for use in developing an acoustic scaling law and testing a first principles trailing edge noise theory. Results were extended to the rotating frame coordinate system to develop a helicopter rotor trailing edge noise prediction. Comparisons of the calculated noise levels with helicopter flyover spectra demonstrate that trailing edge noise contributes significantly to the total helicopter noise spectrum at high frequencies. This noise mechanism is expected to control the minimum rotor noise. In the case of noise radiation from a local blade segment, the acoustic directivity pattern is predicted by the first principles trailing edge noise theory. Acoustic spectra are predicted by a scaling law which includes Mach number, boundary layer thickness and observer position. Spectrum shape and sound pressure level are also predicted by the first principles theory but the analysis does not predict the Strouhal value identifying the spectrum peak.

Lightning protection of the Fokker 100 CFRP rudder

NASA Technical Reports Server (NTRS)

Ruiter, A. J. M.

1991-01-01

The construction of the structural parts of the Fokker 100 CFRP rudder is described with respect to the requirements for electrical bonding and lightning protection. Furthermore, the philosophy for the selection of a consumable trailing edge is given. A description of possible alternative designs for trailing edges and their advantages and disadvantages with respect to damage after lightning impact will also be reviewed. An overview of the tests performed on test samples and the rudder construction are presented and discussed. The effectiveness of both the selected structural provisions and trailing edge are described (and proven) by reporting the results of the simulated lightning tests performed. Proof is given that the trailing edge construction and its bonding through the structural parts of the rudder to the main aircraft structure is a solution which results in minor damage to the rudder after lightning impact. Furthermore, it is shown that the selected trailing edge construction is less favored by the structural designers due to the weight penalty.

The Prediction of Noise Due to Jet Turbulence Convecting Past Flight Vehicle Trailing Edges

NASA Technical Reports Server (NTRS)

Miller, Steven A. E.

2014-01-01

High intensity acoustic radiation occurs when turbulence convects past airframe trailing edges. A mathematical model is developed to predict this acoustic radiation. The model is dependent on the local flow and turbulent statistics above the trailing edge of the flight vehicle airframe. These quantities are dependent on the jet and flight vehicle Mach numbers and jet temperature. A term in the model approximates the turbulent statistics of single-stream heated jet flows and is developed based upon measurement. The developed model is valid for a wide range of jet Mach numbers, jet temperature ratios, and flight vehicle Mach numbers. The model predicts traditional trailing edge noise if the jet is not interacting with the airframe. Predictions of mean-flow quantities and the cross-spectrum of static pressure near the airframe trailing edge are compared with measurement. Finally, predictions of acoustic intensity are compared with measurement and the model is shown to accurately capture the phenomenon.

Acoustic characteristics of externally blown flap systems with mixer nozzles

NASA Technical Reports Server (NTRS)

Goodykoontz, J. H.; Dorsch, R. G.; Wagner, J. M.

1974-01-01

Noise tests were conducted on a large scale, cold flow model of an engine-under-the-wing externally blown flap lift augmentation system employing a mixer nozzle. The mixer nozzle was used to reduce the flap impingement velocity and, consequently, try to attenuate the additional noise caused by the interaction between the jet exhaust and the wing flap. Results from the mixer nozzle tests are summarized and compared with the results for a conical nozzle. The comparison showed that with the mixer nozzle, less noise was generated when the trailing flap was in a typical landing setting (e.g., 60 deg). However, for a takeoff flap setting (20 deg), there was little or no difference in the acoustic characteristics when either the mixer or conical nozzle was used.

Effects of deflected thrust on the longitudinal aerodynamic characteristics of a close-coupled wing-canard configuration. [in the Langley V/STOL tunnel

NASA Technical Reports Server (NTRS)

Yip, L. P.; Paulson, J. W., Jr.

1977-01-01

The effects of power on the longitudinal aerodynamic characteristics of a close-coupled wing-canard fighter configuration with partial-span rectangular nozzles at the trailing edge of the wing were investigated. Data were obtained on a basic wing-strake configuration for nozzle and flap deflections from 0 deg to 30 deg and for nominal thrust coefficients from 0 to 0.30. The model was tested over an angle-of-attack range from -2 deg to 40 deg at Mach numbers of 0.15 and 0.18. Results show substantial improvements in lift-curve slope, in maximum lift, and in drag-due-to-lift efficiency when the canard and strakes have been added to the basic wing-fuselage (wing-alone) configuration. Addition of power increased both lift-curve slope and maximum lift, improved longitudinal stability, and reduced drag due to lift on both the wing-canard and wing-canard-strake configurations. These beneficial effects are primarily derived from boundary-layer control due to moderate thrust coefficients which delay flow separation on the nozzle and inboard portion of the wing flaps.

A Wind-Tunnel Investigation of a Transonic-Transport Configuration Utilizing Drag-Reducing Devices at Mach Numbers from 0.20 to 1.03

NASA Technical Reports Server (NTRS)

Loving, Donald L.

1961-01-01

The static longitudinal stability and control and lateral characteristics of a transonic-transport model, incorporating recent drag-reducing devices, has been investigated in the Langley 8-foot transonic pressure tunnel. The wing was cambered, had a thickened root and a taper ratio of 0.3. Wing sweepback angles of 45 degrees and 40 degrees were investigated with corresponding aspect ratios of 7 and 8, respectively. Modifications to the model for reducing the drag were: a forward fuselage addition and special bodies (four big enough to house jet engines) added to the upper surface of the wing. Other components and changes investigated included an empennage, a wing-tip body, wing fences, wing trailing-edge flaps, horizontal-tail settings, and wing dihedral angle. The investigation covered the Mach number range from 0.20 to 1.03 for the angle-of-attack range from -5 degrees to 15.4 degrees, and a sideslip angle of -5 degrees, in the Reynolds number range from 0.52 times 10(exp 6) to 1.94 times 10(exp 6) based on the wing mean aerodynamic chord. The various fuselage and wing additions delayed the drag-rise Mach number and greatly reduced the drag beyond the drag rise. The wing bodies markedly alleviated unstable pitch tendencies throughout the test Mach number range. At low landing speeds, the wing bodies exhibited little interference with the ability of trailing-edge flaps to increase the lift near maximum lift coefficient; and the use of fences greatly reduced the severe longitudinal instability trend at landing attitudes. The model with a 6 degree dihedral angle exhibited positive lateral and directional stability characteristics in the presence of the fuselage and wing additions. An increase in drag-rise Mach number associated with the fuselage and wing additions on the 40 degree sweptback wing combination was similar to that for the comparable 45 degree combination. These additions did, however, reduce the drag of the 40 degree sweptback configurations more than the 45 degree configurations in the transonic speed range.

Wall Modeled Large Eddy Simulation of Airfoil Trailing Edge Noise

NASA Astrophysics Data System (ADS)

Kocheemoolayil, Joseph; Lele, Sanjiva

2014-11-01

Large eddy simulation (LES) of airfoil trailing edge noise has largely been restricted to low Reynolds numbers due to prohibitive computational cost. Wall modeled LES (WMLES) is a computationally cheaper alternative that makes full-scale Reynolds numbers relevant to large wind turbines accessible. A systematic investigation of trailing edge noise prediction using WMLES is conducted. Detailed comparisons are made with experimental data. The stress boundary condition from a wall model does not constrain the fluctuating velocity to vanish at the wall. This limitation has profound implications for trailing edge noise prediction. The simulation over-predicts the intensity of fluctuating wall pressure and far-field noise. An improved wall model formulation that minimizes the over-prediction of fluctuating wall pressure is proposed and carefully validated. The flow configurations chosen for the study are from the workshop on benchmark problems for airframe noise computations. The large eddy simulation database is used to examine the adequacy of scaling laws that quantify the dependence of trailing edge noise on Mach number, Reynolds number and angle of attack. Simplifying assumptions invoked in engineering approaches towards predicting trailing edge noise are critically evaluated. We gratefully acknowledge financial support from GE Global Research and thank Cascade Technologies Inc. for providing access to their massively-parallel large eddy simulation framework.

Blowing momentum and duty cycle effect on aerodynamic performance of flap by pulsed blowing

NASA Astrophysics Data System (ADS)

Zhou, Ping; Wang, Yankui; Wang, Jinjun; Sha, Yongxiang

2017-04-01

Control surface, which is often located in the trailing edge of wings, is important in the attitude control of an aircraft. However, the efficiency of the control surface declines severely under the high deflect angle of the control surface because of the flow separation. To improve the efficiency of control surface, this study discusses a flow-control technique aimed at suppressing the flow separation by pulsed blowing at the leading edge of the control surface. Results indicated that flow separation over the control surface can be suppressed by pulsed blowing, and the maximum average lift coefficient of the control surface can be 95% times higher than that of without blowing when average blowing momentum coefficient is 0.03 relative to that of without blowing. Finally, this study shows that the average blowing momentum coefficient and non-dimensional frequency of pulsed blowing are two of the key parameters of the pulsed blowing control technique. Otherwise, duty cycle also has influence on the effect of pulsed blowing. Numerical simulation is used in this study.

Unsteady characteristics of a slat-cove flow field

NASA Astrophysics Data System (ADS)

Pascioni, Kyle A.; Cattafesta, Louis N.

2018-03-01

The leading-edge slat of a multielement wing is a significant contributor to the acoustic signature of an aircraft during the approach phase of the flight path. An experimental study of the two-dimensional 30P30N geometry is undertaken to further understand the flow physics and specific noise source mechanisms. The mean statistics from particle image velocimetry (PIV) shows the differences in the flow field with angle of attack, including the interaction between the cove and trailing-edge flow. Phase-locked PIV successfully links narrow-band peaks found in the surface pressure spectrum to shear layer instabilities and also reveals that a bulk cove oscillation at a Strouhal number based on a slat chord of 0.15 exists, indicative of shear layer flapping. Unsteady surface pressure measurements are documented and used to estimate spanwise coherence length scales. A narrow-band frequency prediction scheme is also tested and found to agree well with the data. Furthermore, higher-order spectral analysis suggests that nonlinear effects cause additional peaks to arise in the power spectrum, particularly at low angles of attack.

Unsteady lift and thrust of a 2D flapping thin airfoil in the presence of additional leading edge vortices

NASA Astrophysics Data System (ADS)

Alaminos-Quesada, Javier; Fernandez-Feria, Ramon

2017-11-01

The effect of leading-edge vortices (LEVs) on the lift, thrust and moment of a two-dimensional heaving and pitching foil is analyzed from the unsteady, linear potential theory. General expressions taking into account the effect of unsteady point vortices interacting with the oscillatory trailing wake are first derived. Then, simplified expressions for the initial stages of the growing LEV on each half-stroke are used to obtain analytical closed expressions for the main contribution of these vortices to the lift, thrust and moment. It is found that, within the linear potential framework and the Brown-Michael model, the LEV contributes to the aerodynamic forces and moment only for combined pitching and heaving motions of the foil, being a relevant contribution for sufficiently large values of the product of the reduced frequency and the amplitude of the heaving and/or pitching motions. The results are compared with available experimental data and numerical simulations. Supported by the Ministerio de Economia y Competitividad of Spain Grants No. DPI2013-40479-P and DPI2016-76151-C2-1-R.

Wind-tunnel test results of airfoil modifications for the EA-6B

NASA Technical Reports Server (NTRS)

Sewall, W. G.; Mcghee, R. J.; Ferris, J. C.

1987-01-01

Wind-tunnel tests have been conducted (to determine the effects on airfoil performance for several airfoil modifications) for the EA-6B Wing Improvement Program. The modifications consist of contour changes to the leading-edge slat and trailing-edge flap to provide a higher low-speed maximum lift with no high-speed cruise-drag penalty. Airfoil sections from the 28- and 76-percent span stations were selected as baseline shapes with the major testing devoted to the inboard airfoil section (28-percent span station). The airfoil modifications increased the low-speed maximum lift coefficient between 20 and 35 percent over test conditions of 3 to 14 million chord Reynolds number and 0.14 to 0.34 Mach number. At the high-speed test conditions of 0.4 to 0.80 Mach number and 10 million chord Reynolds number, the modified airfoils had either matched or had lower drag coefficients for all normal-force coefficients above 0.2 as compared to the baseline airfoil. At normal-force coefficients less than 0.2, the baseline (original) airfoil had lower drag coefficients than any of the modified airfoils.

Passive Porous Treatment for Reducing Flap Side-Edge Noise

NASA Technical Reports Server (NTRS)

Choudhari, Meelan M.; Khorrami, Mehdi R.

2008-01-01

A passive porous treatment has been proposed as a means of suppressing noise generated by the airflow around the side edges of partial-span flaps on airplane wings when the flaps are extended in a high-lift configuration. The treatment proposed here does not incur any aerodynamic penalties and could easily be retrofit to existing airplanes. The treatment could also be applied to reduce noise generated by turbomachinery, including wind turbines. Innovative aspects of the proposed treatment include a minimum treatment area and physics-based procedure for treatment design. The efficacy of the treatment was confirmed during wind-tunnel experiments at NASA Ames, wherein the porous treatment was applied to a minute surface area in the vicinity of a flap edge on a 26-percent model of Boeing 777-200 wing.

Detached Eddy Simulation of Flap Side-Edge Flow

NASA Technical Reports Server (NTRS)

Balakrishnan, Shankar K.; Shariff, Karim R.

2016-01-01

Detached Eddy Simulation (DES) of flap side-edge flow was performed with a wing and half-span flap configuration used in previous experimental and numerical studies. The focus of the study is the unsteady flow features responsible for the production of far-field noise. The simulation was performed at a Reynolds number (based on the main wing chord) of 3.7 million. Reynolds Averaged Navier-Stokes (RANS) simulations were performed as a precursor to the DES. The results of these precursor simulations match previous experimental and RANS results closely. Although the present DES simulations have not reached statistical stationary yet, some unsteady features of the developing flap side-edge flowfield are presented. In the final paper it is expected that statistically stationary results will be presented including comparisons of surface pressure spectra with experimental data.

Aeroacoustic measurements on a NACA 0012 applying the Coherent Particle Velocity method

NASA Astrophysics Data System (ADS)

Plogmann, B.; Würz, W.

2013-07-01

Aeroacoustic measurements on two NACA 0012 airfoil sections with different chord length and sharp trailing edge were conducted at the Laminar Wind Tunnel (LWT) of the University of Stuttgart. The LWT is a closed test section wind tunnel with a very low turbulence level and an acoustically optimized diffusor section allowing for high-quality aerodynamic as well as aeroacoustic measurements. Trailing edge noise measurements were performed using the Coherent Particle Velocity (CPV) method, which is based on a cross-spectral analysis of two hot-wire sensor signals placed on the suction and the pressure side of the airfoil trailing edge, respectively. At high angles of attack, the cross-spectral analysis of the two sensor signals used for the measurement of the trailing edge noise can be prone to a disturbing influence of hydrodynamic fluctuations. Hence, continuous shifts in the phasing of the cross-correlation are observed mainly for low sensor distances to the trailing edge. The quantitative evaluation of the trailing edge noise predominately in the low frequency range is, therefore, considerably disturbed. A new approach is proposed, which allows for the correction of the cross-correlation function based on the averaged single wire auto-spectrum. The results are compared to measurements with increased sensor distance and show good agreement. In the following, trailing edge noise measurements were performed on a NACA 0012 airfoil in a wide range of angles of attack ( α = 0°-8°) and free-stream velocities (u_{infty} = 30{-}70 {{m/s}}). The tripped flow cases exhibit a very good consistency for the scaling of the 1/3 octave spectra based on outer variables. Moreover, a common intersection point of the sound pressure level was observed for trailing edge noise spectra measured at constant free-stream velocity and different angles of attack. In cases without boundary layer tripping, the presence of an acoustic feedback loop was observed and linked to the presence of a laminar separation bubble on the pressure side in the vicinity of the trailing edge. Finally, a comparison of the aeroacoustic measurements based on the CPV method showed reasonably good agreement with published data obtained with both a microphone array and the Coherent Output Power method in open-test section facilities.

78 FR 39633 - Airworthiness Directives; The Boeing Company Airplanes

Federal Register 2010, 2011, 2012, 2013, 2014

2013-07-02

...We propose to supersede three existing airworthiness directives (ADs) that apply to The Boeing Company Model 757-200, - 200PF, and -200CB series airplanes. The existing ADs currently require repetitive inspections and audible tap tests of the upper and lower skins of the trailing edge wedges on certain slats, and related investigative and corrective actions if necessary. Since we issued these ADs, we have received reports of slats disbonding on airplanes on which the terminating actions of the existing ADs were completed and also reports of slats disbonding on airplanes outside of the applicability of the existing ADs. This proposed AD would require a determination of the type of trailing edge wedges of the leading edge slats, repetitive inspections on certain trailing edge wedges for areas of skin-to-core disbonding, and corrective actions if necessary. This proposed AD would also provide an optional terminating action for the repetitive inspections. This AD would revise the applicability of the existing ADs to include additional airplanes. We are proposing this AD to prevent delamination of the trailing edge wedge of the leading edge slats, possible loss of pieces of the trailing edge wedge assembly during flight, reduction of the reduced maneuver and stall margins, and consequent reduced controllability of the airplane.

Numerical study on the power extraction performance of a flapping foil with a flexible tail

NASA Astrophysics Data System (ADS)

Wu, J.; Shu, C.; Zhao, N.; Tian, F.-B.

2015-01-01

The numerical study on the power extraction performance of a flapping foil with a flexible tail is performed in this work. A NACA0015 airfoil is arranged in a two-dimensional laminar flow and imposed with a synchronous harmonic plunge and pitch rotary motion. A flat plate that is attached to the trailing edge of the foil is utilized to model a tail, and so they are viewed as a whole for the purpose of power extraction. In addition, the tail either is rigid or can deform due to the exerted hydrodynamic forces. To implement numerical simulations, an immersed boundary-lattice Boltzmann method is employed. At a Reynolds number of 1100 and the position of the pitching axis at third chord, the influences of the mass and flexibility of the tail as well as the frequency of motion on the power extraction are systematically examined. It is found that compared to the foil with a rigid tail, the efficiency of power extraction for the foil with a deformable tail can be improved. Based on the numerical analysis, it is indicated that the enhanced plunging component of the power extraction, which is caused by the increased lift force, directly contributes to the efficiency improvement. Since a flexible tail with medium and high masses is not beneficial to the efficiency improvement, a flexible tail with low mass together with high flexibility is recommended in the flapping foil based power extraction system.

Energy harvesting for self-powered aerostructure actuation

NASA Astrophysics Data System (ADS)

Bryant, Matthew; Pizzonia, Matthew; Mehallow, Michael; Garcia, Ephrahim

2014-04-01

This paper proposes and experimentally investigates applying piezoelectric energy harvesting devices driven by flow induced vibrations to create self-powered actuation of aerostructure surfaces such as tabs, flaps, spoilers, or morphing devices. Recently, we have investigated flow-induced vibrations and limit cycle oscillations due to aeroelastic flutter phenomena in piezoelectric structures as a mechanism to harvest energy from an ambient fluid flow. We will describe how our experimental investigations in a wind tunnel have demonstrated that this harvested energy can be stored and used on-demand to actuate a control surface such as a trailing edge flap in the airflow. This actuated control surface could take the form of a separate and discrete actuated flap, or could constitute rotating or deflecting the oscillating energy harvester itself to produce a non-zero mean angle of attack. Such a rotation of the energy harvester and the associated change in aerodynamic force is shown to influence the operating wind speed range of the device, its limit cycle oscillation (LCO) amplitude, and its harvested power output; hence creating a coupling between the device's performance as an energy harvester and as a control surface. Finally, the induced changes in the lift, pitching moment, and drag acting on a wing model are quantified and compared for a control surface equipped with an oscillating energy harvester and a traditional, static control surface of the same geometry. The results show that when operated in small amplitude LCO the energy harvester adds negligible aerodynamic drag.

Cooling of gas turbines IX : cooling effects from use of ceramic coatings on water-cooled turbine blades

NASA Technical Reports Server (NTRS)

Brown, W Byron; Livingood, John N B

1948-01-01

The hottest part of a turbine blade is likely to be the trailing portion. When the blades are cooled and when water is used as the coolant, the cooling passages are placed as close as possible to the trailing edge in order to cool this portion. In some cases, however, the trailing portion of the blade is so narrow, for aerodynamic reasons, that water passages cannot be located very near the trailing edge. Because ceramic coatings offer the possibility of protection for the trailing part of such narrow blades, a theoretical study has been made of the cooling effect of a ceramic coating on: (1) the blade-metal temperature when the gas temperature is unchanged, and (2) the gas temperature when the metal temperature is unchanged. Comparison is also made between the changes in the blade or gas temperatures produced by ceramic coatings and the changes produced by moving the cooling passages nearer the trailing edge. This comparison was made to provide a standard for evaluating the gains obtainable with ceramic coatings as compared to those obtainable by constructing the turbine blade in such a manner that water passages could be located very near the trailing edge.

Wind tunnel results for a high-speed, natural laminar-flow airfoil designed for general aviation aircraft

NASA Technical Reports Server (NTRS)

Sewall, William G.; Mcghee, Robert J.; Viken, Jeffery K.; Waggoner, Edgar G.; Walker, Betty S.; Millard, Betty F.

1985-01-01

Two dimensional wind tunnel tests were conducted on a high speed natural laminar flow airfoil in both the Langley 6 x 28 inch Transonic Tunnel and the Langley Low Turbulence Pressure Tunnel. The test conditions consisted of Mach numbers ranging from 0.10 to 0.77 and Reynolds numbers ranging from 3 x 1 million to 11 x 1 million. The airfoil was designed for a lift coefficient of 0.20 at a Mach number of 0.70 and Reynolds number of 11 x 1 million. At these conditions, laminar flow would extend back to 50 percent chord of the upper surface and 70 percent chord of the lower surface. Low speed results were also obtained with a 0.20 chord trailing edge split flap deflected 60 deg.

Analytical study of a free-wing/free-trimmer concept. [for gust alleviation and high lift

NASA Technical Reports Server (NTRS)

Porter, R. F.; Hall, D. W.; Brown, J. H., Jr.; Gregorek, G. M.

1978-01-01

The free-wing/free-trimmer is a NASA-Conceived extension of the free-wing concept intended to permit the use of high-lift flaps. Wing pitching moments are balanced by a smaller, external surface attached by a boom or equivalent structure. The external trimmer is, itself, a miniature free wing, and pitch control of the wing-trimmer assembly is effected through a trailing-edge control tab on the trimmer surface. The longitudinal behavior of representative small free-wing/free-trimmer aircraft was analyzed. Aft-mounted trimmer surfaces are found to be superior to forward trimmers, although the permissible trimmer moment arm is limited, in both cases, by adverse dynamic effects. Aft-trimmer configurations provide excellent gust alleviation and meet fundamental stick-fixed stability criteria while exceeding the lift capabilities of pure free-wing configurations.

Recent Turbulence Model Advances Applied to Multielement Airfoil Computations

NASA Technical Reports Server (NTRS)

Rumsey, Christopher L.; Gatski, Thomas B.

2000-01-01

A one-equation linear turbulence model and a two-equation nonlinear explicit algebraic stress model (EASM) are applied to the flow over a multielement airfoil. The effect of the K-epsilon and K-omega forms of the two-equation model are explored, and the K-epsilon form is shown to be deficient in the wall-bounded regions of adverse pressure gradient flows. A new K-omega form of EASM is introduced. Nonlinear terms present in EASM are shown to improve predictions of turbulent shear stress behind the trailing edge of the main element and near midflap. Curvature corrections are applied to both the one- and two-equation turbulence models and yield only relatively small local differences in the flap region, where the flow field undergoes the greatest curvature. Predictions of maximum lift are essentially unaffected by the turbulence model variations studied.

On the wake flow of asymmetrically beveled trailing edges

NASA Astrophysics Data System (ADS)

Guan, Yaoyi; Pröbsting, Stefan; Stephens, David; Gupta, Abhineet; Morris, Scott C.

2016-05-01

Trailing edge and wake flows are of interest for a wide range of applications. Small changes in the design of asymmetrically beveled or semi-rounded trailing edges can result in significant difference in flow features which are relevant for the aerodynamic performance, flow-induced structural vibration and aerodynamically generated sound. The present study describes in detail the flow field characteristics around a family of asymmetrically beveled trailing edges with an enclosed trailing-edge angle of 25° and variable radius of curvature R. The flow fields over the beveled trailing edges are described using data obtained by particle image velocimetry (PIV) experiments. The flow topology for different trailing edges was found to be strongly dependent on the radius of curvature R, with flow separation occurring further downstream as R increases. This variation in the location of flow separation influences the aerodynamic force coefficients, which were evaluated from the PIV data using a control volume approach. Two-point correlations of the in-plane velocity components are considered to assess the structure in the flow field. The analysis shows large-scale coherent motions in the far wake, which are associated with vortex shedding. The wake thickness parameter yf is confirmed as an appropriate length scale to characterize this large-scale roll-up motion in the wake. The development in the very near wake was found to be critically dependent on R. In addition, high-speed PIV measurements provide insight into the spectral characteristics of the turbulent fluctuations. Based on the time-resolved flow field data, the frequency range associated with the shedding of coherent vortex pairs in the wake is identified. By means of time-correlation of the velocity components, turbulent structures are found to convect from the attached or separated shear layers without distinct separation point into the wake.

Flap Side Edge Liners for Airframe Noise Reduction

NASA Technical Reports Server (NTRS)

Jones, Michael G. (Inventor); Khorrami, Mehdi R. (Inventor); Choudhari, Meelan M. (Inventor); Howerton, Brian M. (Inventor)

2014-01-01

One or more acoustic liners comprising internal chambers or passageways that absorb energy from a noise source on the aircraft are disclosed. The acoustic liners may be positioned at the ends of flaps of an aircraft wing to provide broadband noise absorption and/or dampen the noise producing unsteady flow features, and to reduce the amount of noise generated due to unsteady flow at the inboard and/or outboard end edges of a flap.

Optimization of the poro-serrated trailing edges for airfoil broadband noise reduction.

PubMed

Chong, Tze Pei; Dubois, Elisa

2016-08-01

This paper reports an aeroacoustic investigation of a NACA0012 airfoil with a number of poro-serrated trailing edge devices that contain porous materials of various air flow resistances at the gaps between adjacent members of the serrated-sawtooth trailing edge. The main objective of this work is to determine whether multiple-mechanisms on the broadband noise reduction can co-exist on a poro-serrated trailing edge. When the sawtooth gaps are filled with porous material of low-flow resistivity, the vortex shedding tone at low-frequency could not be completely suppressed at high-velocity, but a reasonably good broadband noise reduction can be achieved at high-frequency. When the sawtooth gaps are filled with porous material of very high-flow resistivity, no vortex shedding tone is present, but the serration effect on the broadband noise reduction becomes less effective. An optimal choice of the flow resistivity for a poro-serrated configuration has been identified, where it can surpass the conventional serrated trailing edge of the same geometry by achieving a further 1.5 dB reduction in the broadband noise while completely suppressing the vortex shedding tone. A weakened turbulent boundary layer noise scattering at the poro-serrated trailing edge is reflected by the lower-turbulence intensity at the near wake centreline across the whole spanwise wavelength of the sawtooth.

Theoretical and experimental study of a new method for prediction of profile drag of airfoil sections

NASA Technical Reports Server (NTRS)

Goradia, S. H.; Lilley, D. E.

1975-01-01

Theoretical and experimental studies are described which were conducted for the purpose of developing a new generalized method for the prediction of profile drag of single component airfoil sections with sharp trailing edges. This method aims at solution for the flow in the wake from the airfoil trailing edge to the large distance in the downstream direction; the profile drag of the given airfoil section can then easily be obtained from the momentum balance once the shape of velocity profile at a large distance from the airfoil trailing edge has been computed. Computer program subroutines have been developed for the computation of the profile drag and flow in the airfoil wake on CDC6600 computer. The required inputs to the computer program consist of free stream conditions and the characteristics of the boundary layers at the airfoil trailing edge or at the point of incipient separation in the neighborhood of airfoil trailing edge. The method described is quite generalized and hence can be extended to the solution of the profile drag for multi-component airfoil sections.

On the Connection Between Flap Side-Edge Noise and Tip Vortex Dynamics

NASA Technical Reports Server (NTRS)

Casalino, D.; Hazir, A.; Fares, E.; Duda, B.; Khorrami, M. R.

2015-01-01

The goal of the present work is to investigate how the dynamics of the vortical flow about the flap side edge of an aircraft determine the acoustic radiation. A validated lattice- Boltzmann CFD solution of the unsteady flow about a detailed business jet configuration in approach conditions is used for the present analysis. Evidence of the connection between the noise generated by several segments of the inboard flap tip and the aerodynamic forces acting on the same segments is given, proving that the noise generation mechanism has a spatially coherent and acoustically compact character on the scale of the flap chord, and that the edge-scattering effects are of secondary importance. Subsequently, evidence of the connection between the kinematics of the tip vortex system and the aerodynamic force is provided. The kinematics of the dual vortex system are investigated via a core detection technique. Emphasis is placed on the mutual induction effects between the two main vortices rolling up from the pressure and suction sides of the flap edge. A simple heuristic formula that relates the far-field noise spectrum and the cross-spectrum of the unsteady vortical positions is developed.

Turbulence Measurements on a Flap-Edge Model

NASA Technical Reports Server (NTRS)

Moriarty, Patrick; Bradshaw, Peter; Cantwell, Brian; Ross, James

1998-01-01

Turbulence measurements have been made on a flap-edge and leading-edge slat model using hot-wire anemometry, and, later, particle image velocimetry. The properties of hot-wire anemometry were studied using facilities at NASA Ames Research Center. Hot-film probes were used because of their durability, but cross-films were limited by non-linear end effects. As a warm-up exercise, hot-film probes were used to measure velocities in the farfield wake of a cylinder with an airfoil in the near-field wake. The airfoil reduced the drag coefficient of the system by 10%. A single-wire hot-film probe was used to measure velocity profiles over the top of a NACA 63(sub 2)-215 Mod. B wing with a Fowler flap and leading,-edge slat. Results showed the size of slat wake was dependent upon the slat deflection angle. Velocity increased through the slat gap with increased deflection. The acoustically modified slat decreased the chance of separation. Measurements were taken at the flap edge with a single hot-film. Trends in the data indicate velocity and turbulence levels increase at the flap edge. The acoustically modified flap modifies the mean flow near the flap edge. Correlations were made between the hot-film signal and the unsteady pressure transducers on the wing which were published in a NASA CDTM. The principles of Particle Image Velocimetry (PIV) were studied at Florida State University. Spectral PIV was used to measure the spectra of a subsonic jet. Measured frequencies were close to the predicted frequency of jet shedding. Spectral PIV will be used to measure the spectra of the slat flow in the second 7 x lO-ft. wind tunnel test. PIV has an advantage that it can measure velocity and spectra of the entire flowfield instantaneously. However, problems arise when trying, to store this massive amount of PIV data. Support for this research has continued through a NASA Graduate Student Program Fellowship which will end in June 1999. The thesis should be completed by this time.

12 CFR Appendix D to Part 229 - Indorsement, Reconverting Bank Identification, and Truncating Bank Identification Standards

Code of Federal Regulations, 2010 CFR

2010-01-01

... the leading edge of the check to 1.5 inches from the trailing edge of the check. 31 31 The leading edge is definded as the right side of the check looking at it from the front. The trailing edge is... on the back of the check between 1.88 and 2.74 inches from the leading edge of the check. The...

Boattail Plates With Non-Rectangular Geometries For Reducing Aerodynamic Base Drag Of A Bluff Body In Ground Effect

DOEpatents

Ortega, Jason M.; Sabari, Kambiz

2006-03-07

An apparatus for reducing the aerodynamic base drag of a bluff body having a leading end, a trailing end, a top surface, opposing left and right side surfaces, and a base surface at the trailing end substantially normal to a longitudinal centerline of the bluff body, with the base surface joined (1) to the left side surface at a left trailing edge, (2) to the right side surface at a right trailing edge, and (3) to the top surface at a top trailing edge. The apparatus includes left and right vertical boattail plates which are orthogonally attached to the base surface of the bluff body and inwardly offset from the left and right trailing edges, respectively. This produces left and right vertical channels which generate, in a flowstream substantially parallel to the longitudinal centerline, respective left and right vertically-aligned vortical structures, with the left and right vertical boattail plates each having a plate width defined by a rear edge of the plate spaced from the base surface. Each plate also has a peak plate width at a location between top and bottom ends of the plate corresponding to a peak vortex of the respective vertically-aligned vortical structures.

Reconstruction of pressure sores with perforator-based propeller flaps.

PubMed

Jakubietz, Rafael G; Jakubietz, Danni F; Zahn, Robert; Schmidt, Karsten; Meffert, Rainer H; Jakubietz, Michael G

2011-03-01

Perforator flaps have been successfully used for reconstruction of pressure sores. Although V-Y advancement flaps approximate debrided wound edges, perforator-based propeller flaps allow rotation of healthy tissue into the defect. Perforator-based propeller flaps were planned in 13 patients. Seven pressure sores were over the sacrum, five over the ischial tuberosity, and one on the tip of the scapula. Three patients were paraplegic, six were bedridden, and five were ambulatory. In three patients, no perforators were found. In 10 patients, propeller flaps were transferred. In two patients, total flap necrosis occurred, which was reconstructed with local advancement flaps. In two cases, a wound dehiscence occurred and had to be revised. One hematoma required evacuation. No further complications were noted. No recurrence at the flap site occurred. Local perforator flaps allow closure of pressure sores without harvesting muscle. The propeller version has the added benefit of transferring tissue from a distant site, avoiding reapproximation of original wound edges. Twisting of the pedicle may cause torsion and venous obstruction. This can be avoided by dissecting a pedicle of at least 3 cm. Propeller flaps are a safe option for soft tissue reconstruction of pressure sores. © Thieme Medical Publishers.

Bio-Inspired Control of Roughness and Trailing Edge Noise

NASA Astrophysics Data System (ADS)

Clark, Ian Andrew

Noise from fluid flow over rough surfaces is an important consideration in the design and performance of certain vehicles with high surface-area-to-perimeter ratios. A new method of noise control based on the anatomy of owls is developed and consists of fabric or fibrous canopies suspended above the surface. The method is tested experimentally and is found to reduce the total far-field noise emitted by the surface. The treatment also is found to reduce the magnitude of pressure fluctuations felt by the underlying surface by up to three orders of magnitude. Experimental investigations into the effects of geometric parameters of the canopies lead to an optimized design which maximizes noise reduction. The results obtained during the canopy experiment inspired a separate new device for the reduction of trailing edge noise. This type of noise is generated by flow past the wing of an aircraft or the blades of a wind turbine, and is a source of annoyance for those in surrounding communities. The newly developed treatment consists of small fins, or "finlets," placed near the trailing edge of an airfoil. The treatment is tested experimentally at near-full-scale conditions and is found to reduce the magnitude of far-field noise by up to 10 dB. Geometric parameters of the finlets are tested to determine the optimal size and spacing of the finlets to maximize noise reduction. Follow-up computational and experimental studies reveal the fluid mechanics behind the noise reduction by showing that the finlets produce a velocity deficit in the flow near the trailing edge and limit the magnitude and spanwise correlation lengthscale of turbulence near the trailing edge, factors which determine the magnitude of far-field noise. In a final experiment, the finlets are applied to a marine propeller and are found to reduce not only trailing edge noise, but also noise caused by the bluntness of the trailing edge. The results of this experiment show the potential usefulness of finlets to reduce noise from rotating systems, such as fans or propellers, as well as from structures which feature blunt trailing edges.

A Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) Determined from Phased Microphone Arrays

NASA Technical Reports Server (NTRS)

Brooks, Thomas F.; Humphreys, William M.

2006-01-01

Current processing of acoustic array data is burdened with considerable uncertainty. This study reports an original methodology that serves to demystify array results, reduce misinterpretation, and accurately quantify position and strength of acoustic sources. Traditional array results represent noise sources that are convolved with array beamform response functions, which depend on array geometry, size (with respect to source position and distributions), and frequency. The Deconvolution Approach for the Mapping of Acoustic Sources (DAMAS) method removes beamforming characteristics from output presentations. A unique linear system of equations accounts for reciprocal influence at different locations over the array survey region. It makes no assumption beyond the traditional processing assumption of statistically independent noise sources. The full rank equations are solved with a new robust iterative method. DAMAS is quantitatively validated using archival data from a variety of prior high-lift airframe component noise studies, including flap edge/cove, trailing edge, leading edge, slat, and calibration sources. Presentations are explicit and straightforward, as the noise radiated from a region of interest is determined by simply summing the mean-squared values over that region. DAMAS can fully replace existing array processing and presentations methodology in most applications. It appears to dramatically increase the value of arrays to the field of experimental acoustics.

Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream

DOEpatents

Myers, R.B.; Yagiela, A.S.

1990-12-25

An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member. 3 figs.

Airfoil lance apparatus for homogeneous humidification and sorbent dispersion in a gas stream

DOEpatents

Myers, Robert B.; Yagiela, Anthony S.

1990-12-25

An apparatus for spraying an atomized mixture into a gas stream comprises a stream line airfoil member having a large radius leading edge and a small radius trailing edge. A nozzle assembly pierces the trailing edge of the airfoil member and is concentrically surrounded by a nacelle which directs shielding gas from the interior of the airfoil member around the nozzle assembly. Flowable medium to be atomized and atomizing gas for atomizing the medium are supplied in concentric conduits to the nozzle. A plurality of nozzles each surrounded by a nacelle are spaced along the trailing edge of the airfoil member.

SIMS chemical analysis of extended impacts on the leading and trailing edges of LDEF experiment AO187-2

NASA Technical Reports Server (NTRS)

Amari, S.; Foote, J.; Simon, Charles G.; Swan, P.; Walker, R. M.; Zinner, E.; Jessberger, E. K.; Lange, G.; Stadermann, F.

1992-01-01

The Long Duration Exposure Facility (LDEF) Experiment AO187-2 consisted of 237 capture cells, 120 on the leading edge and 117 on the trailing edge. Each cell was made of polished Ge plates covered with 2.5 micron thick mylar foil at 200 microns from the Ge. Although all leading edge cells and 105 trailing edge cells had lost their plastic covers during flight, optical and electron microscope examination revealed extended impacts in bare cells from either edge that apparently were produced by high velocity projectiles while the plastic foils were still in place. Detailed optical scanning yielded 53 extended impacts on 100 bare cells from the trailing edge that were selected for SIMS chemical analysis. Lateral multi-element ion probe profiles were obtained on 40 of these impacts. Material that can be attributed to the incoming projectiles was found in all analyzed extended compact features and most seem to be associated with cosmic dust particles. However, LDEF deposits are systematically enriched in the refractory elements Al, Ca, and Ti relative to Mg and Fe when compared to IDP's collected in the stratosphere and to chondritic compositions. These differences are most likely due to elemental fractionation effects during the high velocity impact but real differences between interplanetary particles captured on LDEF and stratospheric IDP's cannot be excluded. Recently we extended our studies to cells from the leading edge and the covered cells from the trailing edge. The 12 covered cells contain 20 extended impact candidates. Ion probe analysis of 3 yielded results similar to those obtained for impacts on the bare cells from the trailing edge. Optical scanning of the bare leading edge cell also reveals many extended impacts (42 on 22 cells scanned to date), demonstrating that the cover foils remained intact at least for some time. However, SIMS analysis showed elements that can reasonably be attributed to micrometeoroids in only 2 out of 11 impacts. Eight impacts have residues dominated by Al and one dominated by Ti, indicating a preponderance of orbital debris in leading edge impacts.

Leading edge flap system for aircraft control augmentation

NASA Technical Reports Server (NTRS)

Rao, D. M. (Inventor)

1984-01-01

Traditional roll control systems such as ailerons, elevons or spoilers are least effective at high angles of attack due to boundary layer separation over the wing. This invention uses independently deployed leading edge flaps on the upper surfaces of vortex stabilized wings to shift the center of lift outboard. A rolling moment is created that is used to control roll in flight at high angles of attack. The effectiveness of the rolling moment increases linearly with angle of attack. No adverse yaw effects are induced. In an alternate mode of operation, both leading edge flaps are deployed together at cruise speeds to create a very effective airbrake without appreciable modification in pitching moment. Little trim change is required.

Near Field Trailing Edge Tone Noise Computation

NASA Technical Reports Server (NTRS)

Loh, Ching Y.

2002-01-01

Blunt trailing edges in a flow often generate tone noise due to wall-jet shear layer and vortex shedding. In this paper, the space-time conservation element (CE/SE) method is employed to numerically study the near-field noise of blunt trailing edges. Two typical cases, namely, flow past a circular cylinder (aeolian noise problem) and flow past a flat plate of finite thickness are considered. The computed frequencies compare well with experimental data. For the aeolian noise problem, comparisons with the results of other numerical approaches are also presented.

Synchronized Schlieren method for vortex shedding in cascade during acoustic resonance

NASA Astrophysics Data System (ADS)

Nagashima, T.; Tanida, Y.

1986-10-01

An evaluation is made of synchronized schlieren optical system methods for the simultaneous visualization of both the acoustic wave and vortex shedding phenomena encountered during acoustic resonance excited by vortex shedding from the trailing edges of cascade blades. Attention is given to the case of parallel flat plate blades in throughflow velocities of up to 100 m/s. The acoustic wavefront is found to appear in the trailing edge region and travel upstream when a pair of vortices of opposite sign are fully developed at the trailing edge.

Low-speed airspeed calibration data for a single-engine research-support aircraft

NASA Technical Reports Server (NTRS)

Holmes, B. J.

1980-01-01

A standard service airspeed system on a single engine research support airplane was calibrated by the trailing anemometer method. The effects of flaps, power, sideslip, and lag were evaluated. The factory supplied airspeed calibrations were not sufficiently accurate for high accuracy flight research applications. The trailing anemometer airspeed calibration was conducted to provide the capability to use the research support airplane to perform pace aircraft airspeed calibrations.

Trailing Edge Noise Prediction Based on a New Acoustic Formulation

NASA Technical Reports Server (NTRS)

Casper, J.; Farassat, F.

2002-01-01

A new analytic result in acoustics called 'Formulation 1B,' proposed by Farassat, is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Williams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experiment. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, using both analytical and experimental data on the airfoil surface. The results are compared to analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

Broadband Trailing Edge Noise Predictions in the Time Domain. Revised

NASA Technical Reports Server (NTRS)

Casper, Jay; Farassat, Fereidoun

2003-01-01

A recently developed analytic result in acoustics, "Formulation 1B," is used to compute broadband trailing edge noise from an unsteady surface pressure distribution on a thin airfoil in the time domain. This formulation is a new solution of the Ffowcs Willliams-Hawkings equation with the loading source term, and has been shown in previous research to provide time domain predictions of broadband noise that are in excellent agreement with experimental results. Furthermore, this formulation lends itself readily to rotating reference frames and statistical analysis of broadband trailing edge noise. Formulation 1B is used to calculate the far field noise radiated from the trailing edge of a NACA 0012 airfoil in low Mach number flows, by using both analytical and experimental data on the airfoil surface. The acoustic predictions are compared with analytical results and experimental measurements that are available in the literature. Good agreement between predictions and measurements is obtained.

A simple concept for covering pressure sores: wound edge-based propeller perforator flap.

PubMed

Kelahmetoglu, Osman; Van Landuyt, Koenraad; Yagmur, Caglayan; Sommeling, Casper E; Keles, Musa K; Tayfur, Volkan; Simsek, Tekin; Demirtas, Yener; Guneren, Ethem

2017-12-01

We present a new surgical modification to allow propeller perforator flaps to cover pressure sores at various locations. We used a propeller perforator flap concept based on the detection of newly formed perforator vessels located 1 cm from the wound margin and stimulated by the chronic inflammation process. Between January 2009 and January 2017, 33 wound edge-based propeller perforator flaps were used to cover pressure sores at various locations in 28 patients. In four cases more than one flap was used on the same patient. The patients comprised 18 males and 10 females with a mean age of 41·25 (range, 16-70) years. All patients underwent follow-up for 0-12 months. The mean follow-up duration was 5·03 months. Venous congestion was observed in three flaps that were rotated by 180° (9·1%). However, there was a significant difference between flaps rotated by 90° and 180° according to the complication rate (P = 0·034). Out of 33 flaps, 29 flaps healed uneventfully. Patients were able to sit and lie on their flaps three weeks after surgery. In our study, we were able to obtain satisfying final results using these novel flaps. © 2017 Medicalhelplines.com Inc and John Wiley & Sons Ltd.

Steady-state and transitional aerodynamic characteristics of a wing in simulated heavy rain

NASA Technical Reports Server (NTRS)

Campbell, Bryan A.; Bezos, Gaudy M.

1989-01-01

The steady-state and transient effects of simulated heavy rain on the subsonic aerodynamic characteristics of a wing model were determined in the Langley 14- by 22-Foot Subsonic Tunnel. The 1.29 foot chord wing was comprised of a NACA 23015 airfoil and had an aspect ratio of 6.10. Data were obtained while test variables of liquid water content, angle of attack, and trailing edge flap angle were parametrically varied at dynamic pressures of 10, 30, and 50 psf (i.e., Reynolds numbers of .76x10(6), 1.31x10(6), and 1.69x10(6)). The experimental results showed reductions in lift and increases in drag when in the simulated rain environment. Accompanying this was a reduction of the stall angle of attack by approximately 4 deg. The transient aerodynamic performance during transition from dry to wet steady-state conditions varied between a linear and a nonlinear transition.

KSC-98pc930

NASA Image and Video Library

1998-08-10

In the Tile Fabrication Shop, Tony Rollins, with United Space Alliance, cuts a High-Temperature Reusable Surface Insulation (HRSI) tile on a gun stock contour milling machine. About 70 percent of a Space Shuttle orbiter’s external surface is shielded from heat by a network of more than 24,000 tiles formed from a silica fiber compound. HRSI tiles cover the lower surface of the orbiter, areas around the forward windows, upper body flap, the base heat shield, the "eyeballs" on the front of the Orbital Maneuvering System (OMS) pods, and the leading and trailing edges of the vertical stabilizer and the rudder speed brake. They are generally 6 inches square, but may also be as large as 12 inches square in some areas, and 1 to 5 inches thick. More advanced materials such as Flexible Insulation Blankets have replaced tiles on some upper surfaces of the orbiter

Flight investigation of an underwing nacelle installation of an auxiliary-inlet ejector nozzle with a clamshell flow diverter from Mach 0.6 to 1.3

NASA Technical Reports Server (NTRS)

Head, V. L.

1972-01-01

A nozzle installation of general interest is a podded engine mounted near the aft lower surface of the wing. The effect of this installation on the performance of an auxiliary-inlet ejector nozzle with a clamshell flow diverter was investigated over a Mach number range of 0.6 to 1.3 by using a modified F-106B aircraft. The clamshell flow diverter was tested in a 17 deg position with double-hinged synchronized floating doors. The ejector nozzle trailing-edge flaps were simulated in the closed position with a rigid structure which provided a boattail angle of 10 deg. Primary nozzle area was varied as exhaust gas temperature was varied between 975 and 1561 K. With the nozzle in a subsonic cruise position, the nozzle gross thrust coefficient was 0.918 at a flight Mach number of 0.9.

Tony Rollins prepares a new tile for the Space Shuttle orbiter

NASA Technical Reports Server (NTRS)

1998-01-01

In the Tile Fabrication Shop, Tony Rollins, with United Space Alliance, cuts a High-Temperature Reusable Surface Insulation (HRSI) tile on a gun stock contour milling machine. About 70 percent of a Space Shuttle orbiter's external surface is shielded from heat by a network of more than 24,000 tiles formed from a silica fiber compound. HRSI tiles cover the lower surface of the orbiter, areas around the forward windows, upper body flap, the base heat shield, the 'eyeballs' on the front of the Orbital Maneuvering System (OMS) pods, and the leading and trailing edges of the vertical stabilizer and the rudder speed brake. They are generally 6 inches square, but may also be as large as 12 inches square in some areas, and 1 to 5 inches thick. More advanced materials such as Flexible Insulation Blankets have replaced tiles on some upper surfaces of the orbiter.

Wind-Tunnel Investigation of a Rectangular NACA 2212 Airfoil with Semispan Ailerons and with Nonperforated, Balanced Double Split Flaps for Use as Aerodynamic Brakes

NASA Technical Reports Server (NTRS)

Ivey, Margaret F

1945-01-01

Flat-plate flaps with no wing cutouts and flaps having Clark Y sections with corresponding cutouts made in wing were tested for various flap deflections, chord-wise locations, and gaps between flaps and airfoil contour. The drag was slightly lower for wing with airfoil section flaps. Satisfactory aileron effectiveness was obtained with flap gap of 20% wing chord and flap-nose location of 80 percent wing chord behind leading edge. Airflow was smooth and buffeting negligible.

Unsteady aerodynamics and vortex-sheet formation of a two-dimensional airfoil

NASA Astrophysics Data System (ADS)

Xia, X.; Mohseni, K.

2017-11-01

Unsteady inviscid flow models of wings and airfoils have been developed to study the aerodynamics of natural and man-made flyers. Vortex methods have been extensively applied to reduce the dimensionality of these aerodynamic models, based on the proper estimation of the strength and distribution of the vortices in the wake. In such modeling approaches, one of the most fundamental questions is how the vortex sheets are generated and released from sharp edges. To determine the formation of the trailing-edge vortex sheet, the classical Kutta condition can be extended to unsteady situations by realizing that a flow cannot turn abruptly around a sharp edge. This condition can be readily applied to a flat plate or an airfoil with cusped trailing edge since the direction of the forming vortex sheet is known to be tangential to the trailing edge. However, for a finite-angle trailing edge, or in the case of flow separation away from a sharp corner, the direction of the forming vortex sheet is ambiguous. To remove any ad-hoc implementation, the unsteady Kutta condition, the conservation of circulation, as well as the conservation laws of mass and momentum are coupled to analytically solve for the angle, strength, and relative velocity of the trailing-edge vortex sheet. The two-dimensional aerodynamic model together with the proposed vortex-sheet formation condition is verified by comparing flow structures and force calculations with experimental results for airfoils in steady and unsteady background flows.

Edge of polar cap patches

NASA Astrophysics Data System (ADS)

Hosokawa, K.; Taguchi, S.; Ogawa, Y.

2016-04-01

On the night of 4 December 2013, a sequence of polar cap patches was captured by an all-sky airglow imager (ASI) in Longyearbyen, Norway (78.1°N, 15.5°E). The 630.0 nm airglow images from the ASI of 4 second exposure time, oversampled the emission of natural lifetime (with quenching) of at least ˜30 sec, introduce no observational blurring effects. By using such high-quality ASI images, we succeeded in visualizing an asymmetry in the gradients between the leading/trailing edges of the patches in a 2-D fashion. The gradient in the leading edge was found to be 2-3 times steeper than that in the trailing edge. We also identified fingerlike structures, appearing only along the trailing edge of the patches, whose horizontal scale size ranged from 55 to 210 km. These fingers are considered to be manifestations of plasma structuring through the gradient-drift instability (GDI), which is known to occur only along the trailing edge of patches. That is, the current 2-D observations visualized, for the first time, how GDI stirs the patch plasma and such a mixing process makes the trailing edge more gradual. This result strongly implies a close connection between the GDI-driven plasma stirring and the asymmetry in the large-scale shape of patches and then suggests that the fingerlike structures can be used as markers to estimate the fine-scale structure in the plasma flow within patches.

Analysis of Low-Speed Stall Aerodynamics of a Swept Wing with Seamless Flaps

NASA Technical Reports Server (NTRS)

Bui, Trong T.

2016-01-01

Computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a Gulfstream G-III airplane (Gulfstream Aerospace Corporation, Savannah, Georgia) swept wing modified with an experimental seamless, compliant flap called the Adaptive Compliant Trailing Edge (ACTE) flap. The stall characteristics of the modified ACTE wing were analyzed and compared with the unmodified, clean wing at the flight speed of 120 knots and altitude of 2300 feet above mean sea level, in free air as well as in ground effect. A polyhedral finite-volume unstructured full Navier-Stokes CFD code, STAR-CCM (registered trademark) plus (CD-adapco [Computational Dynamics Limited, United Kingdom, and Analysis & Design Application Co., United States]), was used. Steady Reynolds-averaged Navier-Stokes CFD simulations were conducted for a clean wing and the ACTE wings at various ACTE deflection angles in free air (-2 degrees, 15 degrees, and 30 degrees) as well as in ground effect (15 degrees and 30 degrees). Solution sensitivities to grid densities were examined. In free air, the ACTE wings are predicted to stall at lower angles of attack than the clean wing. In ground effect, all wings are predicted to stall at lower angles of attack than the corresponding wings in free air. Even though the lift curves are higher in ground effect than in free air, the maximum lift coefficients for all wings are lower in ground effect. Finally, the lift increase due to ground effect for the ACTE wing is predicted to be less than the clean wing.

An experimental study of airfoil instability tonal noise with trailing edge serrations

NASA Astrophysics Data System (ADS)

Chong, Tze Pei; Joseph, Phillip F.

2013-11-01

This paper presents an experimental study of the effect of trailing edge serrations on airfoil instability noise. Detailed aeroacoustic measurements are presented of the noise radiated by an NACA-0012 airfoil with trailing edge serrations in a low to moderate speed flow under acoustical free field conditions. The existence of a separated boundary layer near the trailing edge of the airfoil at an angle of attack of 4.2 degree has been experimentally identified by a surface mounted hot-film arrays technique. Hot-wire results have shown that the saw-tooth surface can trigger a bypass transition and prevent the boundary layer from becoming separated. Without the separated boundary layer to act as an amplifier for the incoming Tollmien-Schlichting waves, the intensity and spectral characteristic of the radiated tonal noise can be affected depending upon the serration geometry. Particle Imaging Velocimetry (PIV) measurements of the airfoil wakes for a straight and serrated trailing edge are also reported in this paper. These measurements show that localized normal-component velocity fluctuations that are present in a small region of the wake from the laminar airfoil become weakened once serrations are introduced. Owing to the above unique characteristics of the serrated trailing edges, we are able to further investigate the mechanisms of airfoil instability tonal noise with special emphasis on the assessment of the wake and non-wake based aeroacoustic feedback models. It has been shown that the instability tonal noise generated at an angle of attack below approximately one degree could involve several complex mechanisms. On the other hand, the non-wake based aeroacoustic feedback mechanism alone is sufficient to predict all discrete tone frequencies accurately when the airfoil is at a moderate angle of attack. Larger Δf, which is defined as (fn+1-fn). In other words, a larger margin of velocity increase is required in order to "shift" the fn and fn+1 across fs sequentially, which is the condition for a ladder jump to occur. Lower amplification factor A for the T-S waves, which can result in a radiation of lower noise levels for the broadband hump peak at fs. This phenomenon will proportionally reduce the noise level difference for fn and fn+1, thus making an identification of a ladder jump event more difficult. Finally, we believe that the tone noise generated in this experimental study is of genuine tones of an isolated airfoil. This can be supported by the fact that, when considering either a straight trailing edge or a serrated trailing edge, the overall airfoil geometry at the same angle of attack is still retained and the wind tunnel setup and locations of the laboratory equipment, which could potentially become an anchor point for the acoustic feedback loop, are exactly the same. However, the straight S0 and serrated S2" trailing edges have been shown to produce systematically different spectral characteristics, especially in the forms of tonal rungs, which can be predicted accurately by the original and slightly modified acoustic feedback Model A respectively. In summary, the trailing edge serration is a useful device for the suppression of airfoil instability self-noise. For greater control effectiveness, however, the laminar separation bubble must be situated within the serration region of the trailing edge. This connection imposes restrictions on the angle of attack and velocity over which trailing edge serrations are effective. The feedback loop structure about the wake noise source and the suction surface of the airfoil in Model B is ignored in the present case. This assumption should be reasonably valid given that, in our previous study [3], we cannot identify any significant role of the boundary layer flow at the suction surface in contributing the instability tonal noise radiation across a wide range of Reynolds numbers.

Low transient thermal stress turbine engine components

DOEpatents

Shi, Jun [Glastonbury, CT; Schmidt, Wayde R [Pomfret Center, CT

2011-06-28

A turbine vane includes a platform; and at least one airfoil mounted to the platform and having a trailing edge and a leading edge, wherein the vane is composed of a functionally graded material having a first material and a second material, wherein the trailing edge includes a greater amount of the first material than the second material, and the leading edge includes a greater amount of the second material than the first material.

Rotor blades for turbine engines

DOEpatents

Piersall, Matthew R; Potter, Brian D

2013-02-12

A tip shroud that includes a plurality of damping fins, each damping fin including a substantially non-radially-aligned surface that is configured to make contact with a tip shroud of a neighboring rotor blade. At least one damping fin may include a leading edge damping fin and at least one damping fin may include a trailing edge damping fin. The leading edge damping fin may be configured to correspond to the trailing edge damping fin.

Effects of wing leading-edge flap deflections on subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing

NASA Technical Reports Server (NTRS)

Henderson, W. P.

1978-01-01

An investigation was conducted to determine the effects of wing leading-edge flap deflections on the subsonic longitudinal aerodynamic characteristics of a wing-fuselage configuration with a 44 deg swept wing. The tests were conducted at Mach numbers from 0.40 to 0.85, corresponding to Reynolds numbers (based on wing mean geometric chord) of 2.37 x 1,000,000 to 4.59 x 1,000,000 and at angles of attack from -3 deg to 22 deg. The configurations under study included a wing-fuselage configuration and a wing-fuselage-strake configuration. Each configuration had multisegmented, constant-chord leading-edge flaps which could be deflected independently or in various combinations.

An analytical parametric study of the broadband noise from axial-flow fans

NASA Technical Reports Server (NTRS)

Chou, Shau-Tak; George, Albert R.

1987-01-01

The rotating dipole analysis of Ffowcs Williams and Hawkings (1969) is used to predict the far field noise radiation due to various rotor broadband noise mechanisms. Consideration is given to inflow turbulence noise, attached boundary layer/trailing-edge interaction noise, tip-vortex formation noise, and trailing-edge thickness noise. The parametric dependence of broadband noise from unducted axial-flow fans on several critical variables is studied theoretically. The angle of attack of the rotor blades, which is related to the rotor performance, is shown to be important to the trailing-edge noise and to the tip-vortex formation noise.

Structurally efficient inflatable protective device

DOEpatents

Nelsen, James M.; Whinery, Larry D.; Gwinn, Kenneth W.; McBride, Donald D.; Luna, Daniel A.; Holder, Joseph P.; Bliton, Richard J.

1997-01-01

An apparatus and method for making a low cost, self-venting, inflatable protective cushion of simple and structurally efficient design with a shape and construction that optimizes the cushion's ability to withstand inflation pressures and impact when deployed which includes a sheet defined by at least one fold line and a plurality of flap portions, each flap portion having a base edge corresponding to a fold line and at least two side edges each extending outwardly from a base edge and ultimately converging to meet each other, the flap portions being folded at the fold line(s) and being joined at corresponding side edges to define an inflatable chamber. The inflatable protective cushion and method for making same may further include a lightweight, low permeability, fabric that optimizes the cushion's ability to withstand inflation pressures and impact when deployed and minimizes the packed volume of the cushion when stored.

Structurally efficient inflatable protective device

DOEpatents

Nelsen, J.M.; Whinery, L.D.; Gwinn, K.W.; McBride, D.D.; Luna, D.A.; Holder, J.P.; Bliton, R.J.

1997-03-04

An apparatus and method are disclosed for making a low cost, self-venting, inflatable protective cushion of simple and structurally efficient design with a shape and construction that optimizes the cushion`s ability to withstand inflation pressures and impact when deployed which includes a sheet defined by at least one fold line and a plurality of flap portions, each flap portion having a base edge corresponding to a fold line and at least two side edges each extending outwardly from a base edge and ultimately converging to meet each other, the flap portions being folded at the fold line(s) and being joined at corresponding side edges to define an inflatable chamber. The inflatable protective cushion and method for making same may further include a lightweight, low permeability, fabric that optimizes the cushion`s ability to withstand inflation pressures and impact when deployed and minimizes the packed volume of the cushion when stored. 22 figs.

Structurally efficient inflatable protective device

DOEpatents

Nelsen, J.M.; Whinery, L.D.; Gwinn, K.W.; McBride, D.D.; Luna, D.A.; Holder, J.P.; Bliton, R.J.

1996-01-09

An apparatus and method are disclosed for making a low cost, self-venting, inflatable protective cushion of simple and structurally efficient design with a shape and construction that optimizes the cushion`s ability to withstand inflation pressures and impact when deployed which includes a sheet defined by at least one fold line and a plurality of flap portions, each flap portion having a base edge corresponding to a fold line and at least two side edges each extending outwardly from a base edge and ultimately converging to meet each other, the flap portions being folded at the fold line(s) and being joined at corresponding side edges to define an inflatable chamber. The inflatable protective cushion and method for making same may further include a lightweight, low permeability, fabric that optimizes the cushion`s ability to withstand inflation pressures and impact when deployed and minimizes the packed volume of the cushion when stored. 22 figs.

Structurally efficient inflatable protective device

DOEpatents

Nelsen, James M.; Whinery, Larry D.; Gwinn, Kenneth W.; McBride, Donald D.; Luna, Daniel A.; Holder, Joseph P.; Bliton, Richard J.

1996-01-01

An apparatus and method for making a low cost, self-venting, inflatable protective cushion of simple and structurally efficient design with a shape and construction that optimizes the cushion's ability to withstand inflation pressures and impact when deployed which includes a sheet defined by at least one fold line and a plurality of flap portions, each flap portion having a base edge corresponding to a fold line and at least two side edges each extending outwardly from a base edge and ultimately converging to meet each other, the flap portions being folded at the fold line(s) and being Joined at corresponding side edges to define an inflatable chamber. The inflatable protective cushion and method for making same may further include a lightweight, low permeability, fabric that optimizes the cushion's ability to withstand inflation pressures and impact when deployed and minimizes the packed volume of the cushion when stored.

Structurally efficient inflatable protective device

DOEpatents

Nelsen, James M.; Whinery, Larry D.; Gwinn, Kenneth W.; McBride, Donald D.; Luna, Daniel A.; Holder, Joseph P.; Bliton, Richard J.

1996-01-01

An apparatus and method for making a low cost, self-venting, inflatable protective cushion of simple and structurally efficient design with a shape and construction that optimizes the cushion's ability to withstand inflation pressures and impact when deployed which includes a sheet defined by at least one fold line and a plurality of flap portions, each flap portion having a base edge corresponding to a fold line and at least two side edges each extending outwardly from a base edge and ultimately converging to meet each other, the flap portions being folded at the fold line(s) and being joined at corresponding side edges to define an inflatable chamber. The inflatable protective cushion and method for making same may further include a lightweight, low permeability, fabric that optimizes the cushion's ability to withstand inflation pressures and impact when deployed and minimizes the packed volume of the cushion when stored.

A new approach to complete aircraft landing gear noise prediction

NASA Astrophysics Data System (ADS)

Lopes, Leonard V.

This thesis describes a new landing gear noise prediction system developed at The Pennsylvania State University, called Landing Gear Model and Acoustic Prediction code (LGMAP). LGMAP is used to predict the noise of an isolated or installed landing gear geometry. The predictions include several techniques to approximate the aeroacoustic and aerodynamic interactions of landing gear noise generation. These include (1) a method for approximating the shielding of noise caused by the landing gear geometry, (2) accounting for local flow variations due to the wing geometry, (3) the interaction of the landing gear wake with high-lift devices, and (4) a method for estimating the effect of gross landing gear design changes on local flow and acoustic radiation. The LGMAP aeroacoustic prediction system has been created to predict the noise generated by a given landing gear. The landing gear is modeled as a set of simple components that represent individual parts of the structure. Each component, ranging from large to small, is represented by a simple geometric shape and the unsteady flow on the component is modeled based on an individual characteristic length, local flow velocity, and the turbulent flow environment. A small set of universal models is developed and applied to a large range of similar components. These universal models, combined with the actual component geometry and local environment, give a unique loading spectrum and acoustic field for each component. Then, the sum of all the individual components in the complete configuration is used to model the high level of geometric complexity typical of current aircraft undercarriage designs. A line of sight shielding algorithm based on scattering by a two-dimensional cylinder approximates the effect of acoustic shielding caused by the landing gear. Using the scattering from a cylinder in two-dimensions at an observer position directly behind the cylinder, LGMAP is able to estimate the reduction in noise due to shielding by the landing gear geometry. This thesis compares predictions with data from a recent wind tunnel experiment conducted at NASA Langley Research Center, and demonstrates that including the acoustic scattering can improve the predictions by LGMAP at all observer positions. In this way, LGMAP provides more information about the actual noise propagation than simple empirical schemes. Two-dimensional FLUENT calculations of approximate wing cross-sections are used by LGMAP to compute the change in noise due to the change in local flow velocity in the vicinity of the landing gear due to circulation around the wing. By varying angle of attack and flap deflection angle in the CFD calculations, LGMAP is able to predict the noise level change due to the change in local flow velocity in the landing gear vicinity. A brief trade study is performed on the angle of attack of the wing and flap deflection angle of the flap system. It is shown that increasing the angle of attack or flap deflection angle reduces the flow velocity in the vicinity of the landing gear, and therefore the predicted noise. Predictions demonstrate the ability of the prediction system to quickly estimate the change in landing gear noise caused by a change in wing configuration. A three-dimensional immersed boundary CFD calculation of simplified landing gear geometries provides relatively quick estimates of the mean flow around the landing gear. The mean flow calculation provides the landing gear wake geometry for the prediction of trailing edge noise associated with the interaction of the landing gear wake with the high lift devices. Using wind tunnel experiments that relate turbulent intensity to wake size and the Ffowcs Williams and Hall trailing edge noise equation for the acoustic calculation, LGMAP is able to predict the landing gear wake generated trailing edge noise. In this manner, LGMAP includes the effect of the interaction of the landing gear's wake with the wing/flap system on the radiated noise. The final prediction technique implemented includes local flow calculations of a landing gear with various truck angles using the immersed boundary scheme. Using the mean flow calculation, LGMAP is able to predict noise changes caused by gross changes in landing gear design. Calculations of the mean flow around the landing gear show that the rear wheels of a six-wheel bogie experience significantly reduced mean flow velocity when the truck is placed in a toe-down configuration. This reduction in the mean flow results is a lower noise signature from the rear wheel. Since the noise from a six-wheel bogie at flyover observer positions is primarily composed of wheel noise, the reduced local flow velocity results in a reduced noise signature from the entire landing gear geometry. Comparisons with measurements show the accuracy of the predictions of landing gear noise levels and directivity. Airframe noise predictions for the landing gear of a complete aircraft are described including all of the above mentioned developments and prediction techniques. These show that the nose gear noise and the landing gear wake/flap interaction noise, while not significantly changing the overall shape of the radiated noise, do contribute to the overall noise from the installed landing gear.

Experimental Investigation of a Point Design Optimized Arrow Wing HSCT Configuration

NASA Technical Reports Server (NTRS)

Narducci, Robert P.; Sundaram, P.; Agrawal, Shreekant; Cheung, S.; Arslan, A. E.; Martin, G. L.

1999-01-01

The M2.4-7A Arrow Wing HSCT configuration was optimized for straight and level cruise at a Mach number of 2.4 and a lift coefficient of 0.10. A quasi-Newton optimization scheme maximized the lift-to-drag ratio (by minimizing drag-to-lift) using Euler solutions from FL067 to estimate the lift and drag forces. A 1.675% wind-tunnel model of the Opt5 HSCT configuration was built to validate the design methodology. Experimental data gathered at the NASA Langley Unitary Plan Wind Tunnel (UPWT) section #2 facility verified CFL3D Euler and Navier-Stokes predictions of the Opt5 performance at the design point. In turn, CFL3D confirmed the improvement in the lift-to-drag ratio obtained during the optimization, thus validating the design procedure. A data base at off-design conditions was obtained during three wind-tunnel tests. The entry into NASA Langley UPWT section #2 obtained data at a free stream Mach number, M(sub infinity), of 2.55 as well as the design Mach number, M(sub infinity)=2.4. Data from a Mach number range of 1.8 to 2.4 was taken at UPWT section #1. Transonic and low supersonic Mach numbers, M(sub infinity)=0.6 to 1.2, was gathered at the NASA Langley 16 ft. Transonic Wind Tunnel (TWT). In addition to good agreement between CFD and experimental data, highlights from the wind-tunnel tests include a trip dot study suggesting a linear relationship between trip dot drag and Mach number, an aeroelastic study that measured the outboard wing deflection and twist, and a flap scheduling study that identifies the possibility of only one leading-edge and trailing-edge flap setting for transonic cruise and another for low supersonic acceleration.

Glove box shield

DOEpatents

Brackenbush, L.W.; Hoenes, G.R.

A shield for a glove box housing radioactive material is comprised of spaced apart clamping members which maintain three overlapping flaps in place therebetween. There is a central flap and two side flaps, the side flaps overlapping at the interior edges thereof and the central flap extending past the intersection of the side flaps in order to insure that the shield is always closed when the user wthdraws his hand from the glove box. Lead loaded neoprene rubber is the preferred material for the three flaps, the extent of lead loading depending upon the radiation levels within the glove box.

Numerical Calculations of 3-D High-Lift Flows and Comparison with Experiment

NASA Technical Reports Server (NTRS)

Compton, William B, III

2015-01-01

Solutions were obtained with the Navier-Stokes CFD code TLNS3D to predict the flow about the NASA Trapezoidal Wing, a high-lift wing composed of three elements: the main-wing element, a deployed leading-edge slat, and a deployed trailing-edge flap. Turbulence was modeled by the Spalart-Allmaras one-equation turbulence model. One case with massive separation was repeated using Menter's two-equation SST (Menter's Shear Stress Transport) k-omega turbulence model in an attempt to improve the agreement with experiment. The investigation was conducted at a free stream Mach number of 0.2, and at angles of attack ranging from 10.004 degrees to 34.858 degrees. The Reynolds number based on the mean aerodynamic chord of the wing was 4.3 x 10 (sup 6). Compared to experiment, the numerical procedure predicted the surface pressures very well at angles of attack in the linear range of the lift. However, computed maximum lift was 5% low. Drag was mainly under predicted. The procedure correctly predicted several well-known trends and features of high-lift flows, such as off-body separation. The two turbulence models yielded significantly different solutions for the repeated case.

Modeling and Prediction of Krueger Device Noise

NASA Technical Reports Server (NTRS)

Guo, Yueping; Burley, Casey L.; Thomas, Russell H.

2016-01-01

This paper presents the development of a noise prediction model for aircraft Krueger flap devices that are considered as alternatives to leading edge slotted slats. The prediction model decomposes the total Krueger noise into four components, generated by the unsteady flows, respectively, in the cove under the pressure side surface of the Krueger, in the gap between the Krueger trailing edge and the main wing, around the brackets supporting the Krueger device, and around the cavity on the lower side of the main wing. For each noise component, the modeling follows a physics-based approach that aims at capturing the dominant noise-generating features in the flow and developing correlations between the noise and the flow parameters that control the noise generation processes. The far field noise is modeled using each of the four noise component's respective spectral functions, far field directivities, Mach number dependencies, component amplitudes, and other parametric trends. Preliminary validations are carried out by using small scale experimental data, and two applications are discussed; one for conventional aircraft and the other for advanced configurations. The former focuses on the parametric trends of Krueger noise on design parameters, while the latter reveals its importance in relation to other airframe noise components.

Prediction of effects of wing contour modifications on low-speed maximum lift and transonic performance for the EA-6B aircraft

NASA Technical Reports Server (NTRS)

Allison, Dennis O.; Waggoner, E. G.

1990-01-01

Computational predictions of the effects of wing contour modifications on maximum lift and transonic performance were made and verified against low speed and transonic wind tunnel data. This effort was part of a program to improve the maneuvering capability of the EA-6B electronics countermeasures aircraft, which evolved from the A-6 attack aircraft. The predictions were based on results from three computer codes which all include viscous effects: MCARF, a 2-D subsonic panel code; TAWFIVE, a transonic full potential code; and WBPPW, a transonic small disturbance potential flow code. The modifications were previously designed with the aid of these and other codes. The wing modifications consists of contour changes to the leading edge slats and trailing edge flaps and were designed for increased maximum lift with minimum effect on transonic performance. The prediction of the effects of the modifications are presented, with emphasis on verification through comparisons with wind tunnel data from the National Transonic Facility. Attention is focused on increments in low speed maximum lift and increments in transonic lift, pitching moment, and drag resulting from the contour modifications.

SIMS chemical analysis of extended impacts on the leading and trailing edges of LDEF experiment AO187-2

NASA Technical Reports Server (NTRS)

Amari, S.; Foote, J.; Swan, P.; Walker, R. M.; Zinner, E.; Lange, G.

1993-01-01

Numerous 'extended impacts' found in both leading and trailing edge capture cells were successfully analyzed for the chemical composition of projectile residues by secondary ion mass spectrometry (SIMS). Most data were obtained from the trailing edge cells where 45 of 58 impacts were classified as 'probably natural' and the remainder as 'possibly man-made debris.' This is in striking contrast to leading edge cells where 9 of 11 impacts so far measured are definitely classified as orbital debris. Although all the leading edge cells had lost their plastic entrance foils during flight, the rate of foil failure was similar to that of the trailing edge cells, 10 percent of which were recovered intact. Ultraviolet embrittlement is suspected as the major cause of failure on both leading and trailing edges. The major impediment to the accurate determination of projectile chemistry is the fractionation of volatile and refractory elements in the hypervelocity impact and redeposition processes. This effect had been noted in a simulation experiment but is more pronounced in the LDEF capture cells, probably due to the higher average velocities of the space impacts. Surface contamination of the pure Ge surfaces with a substance rich in Si, but also containing Mg and Al, provides an additional problem for the accurate determination of impactor chemistry. The effect is variable, being much larger on surfaces that were exposed to space than in those cells that remained intact. Future work will concentrate on the analyses of more leading edge impacts and the development of new SIMS techniques for the measurement of elemental abundances in extended impacts.

Influence of wing tip morphology on vortex dynamics of flapping flight

NASA Astrophysics Data System (ADS)

Krishna, Swathi; Mulleners, Karen

2013-11-01

The mechanism of flapping wing flight provides insects with extraordinary flight capabilities. The uniquely shaped wing tips give insects an edge in flight performance and the interaction between the leading edge vortices and wing tip vortices enhance their propelling efficiencies and manoeuvrability. These are qualities that are sought after in current-day Micro Air Vehicles. A detailed understanding of the vortex dynamics of flapping flight and the influence of the wing tip planform is imperative for technical application. An experimental study is conducted to investigate the effects of different wing tip planforms on the formation, evolution and interaction of vortical structures. We thereby focus on the interaction between the coherent structures evolving from the leading edge and the wing tip during pitching and flapping motions.The spatial and temporal evolution of the three-dimensional flow structures are determined using Scanning (Stereo) Particle Image Velocimetry and an in-depth coherent structure analysis. By comparing the vortex dynamics, the aerodynamic performance of various wing tip planforms are evaluated.

A novel capsulorhexis technique using shearing forces with cystotome.

PubMed

Karim, Shah M R; Ong, Chin T; Sleep, Tamsin J

2010-05-15

To demonstrate a capsulorhexis technique using predominantly shearing forces with a cystotome on a virtual reality simulator and on a human eye. Our technique involves creating the initial anterior capsular tear with a cystotome to raise a flap. The flap left unfolded on the lens surface. The cystotome tip is tilted horizontally and is engaged on the flap near the leading edge of the tear. The cystotome is moved in a circular fashion to direct the vector forces. The loose flap is constantly swept towards the centre so that it does not obscure the view on the tearing edge. Our technique has the advantage of reducing corneal wound distortion and subsequent anterior chamber collapse. The capsulorhexis flap is moved away from the tear leading edge allowing better visualisation of the direction of tear. This technique offers superior control of the capsulorhexis by allowing the surgeon to change the direction of the tear to achieve the desired capsulorhexis size. The EYESI Surgical Simulator is a realistic training platform for surgeons to practice complex capsulorhexis techniques. The shearing forces technique is a suitable alternative and in some cases a far better technique in achieving the desired capsulorhexis.

On the Importance of Spatial Resolution for Flap Side Edge Noise Prediction

NASA Technical Reports Server (NTRS)

Mineck, Raymond E.; Khorrami, Mehdi R.

2017-01-01

A spatial resolution study of flap tip flow and the effects on the farfield noise signature for an 18%-scale, semispan Gulfstream aircraft model are presented. The NASA FUN3D unstructured, compressible Navier-Stokes solver was used to perform the highly resolved, time-dependent, detached eddy simulations of the flow field associated with the flap for this high-fidelity aircraft model. Following our previous work on the same model, the latest computations were undertaken to determine the causes of deficiencies observed in our earlier predictions of the steady and unsteady surface pressures and off-surface flow field at the flap tip regions, in particular the outboard tip area, where the presence of a cavity at the side-edge produces very complex flow features and interactions. The present results show gradual improvement in steady loading at the outboard flap edge region with increasing spatial resolution, yielding more accurate fluctuating surface pressures, off-surface flow field, and farfield noise with improved high-frequency content when compared with wind tunnel measurements. The spatial resolution trends observed in the present study demonstrate that the deficiencies reported in our previous computations are mostly caused by inadequate spatial resolution and are not related to the turbulence model.

Vortex developments over steady and accelerated airfoils incorporating a trailing edge jet

NASA Technical Reports Server (NTRS)

Finaish, F.; Okong'o, N.; Frigerio, J.

1993-01-01

Computational and experimental studies are conducted to investigate the influence of a trailing edge jet on flow separation and subsequent vortex formation over steady and accelerated airfoils at high angles of attack. A computer code, employing the stream function-vorticity approach, is developed and utilized to conduct numerical experiments on the flow problem. To verify and economize such efforts, an experimental system is developed and incorporated into a subsonic wind tunnel where streamline and vortex flow visualization experiments are conducted. The study demonstrates the role of the trailing edge jet in controlling flow separation and subsequent vortex development for steady and accelerating flow at angles past the static stall angle of attack. The results suggest that the concept of the trailing edge jet may be utilized to control the characteristics of unsteady separated flows over lifting surfaces. This control possibility seems to be quite effective and could have a significant role in controlling unsteady separated flows.

Numerical design of advanced multi-element airfoils

NASA Technical Reports Server (NTRS)

Mathias, Donovan L.; Cummings, Russell M.

1994-01-01

The current study extends the application of computational fluid dynamics to three-dimensional high-lift systems. Structured, overset grids are used in conjunction with an incompressible Navier-Stokes flow solver to investigate flow over a two-element high-lift configuration. The computations were run in a fully turbulent mode using the one-equation Baldwin-Barth turbulence model. The geometry consisted of an unswept wing which spanned a wind tunnel test section. Flows over full and half-span Fowler flap configurations were computed. Grid resolution issues were investigated in two dimensional studies of the flapped airfoil. Results of the full-span flap wing agreed well with experimental data and verified the method. Flow over the wing with the half-span was computed to investigate the details of the flow at the free edge of the flap. The results illustrated changes in flow streamlines, separation locations, and surface pressures due to the vortex shed from the flap edge.

VORCOR: A computer program for calculating characteristics of wings with edge vortex separation by using a vortex-filament and-core model

NASA Technical Reports Server (NTRS)

Pao, J. L.; Mehrotra, S. C.; Lan, C. E.

1982-01-01

A computer code base on an improved vortex filament/vortex core method for predicting aerodynamic characteristics of slender wings with edge vortex separations is developed. The code is applicable to camber wings, straked wings or wings with leading edge vortex flaps at subsonic speeds. The prediction of lifting pressure distribution and the computer time are improved by using a pair of concentrated vortex cores above the wing surface. The main features of this computer program are: (1) arbitrary camber shape may be defined and an option for exactly defining leading edge flap geometry is also provided; (2) the side edge vortex system is incorporated.

Implementation of a Biaxial Resonant Fatigue Test Method on a Large Wind Turbine Blade

DOE Office of Scientific and Technical Information (OSTI.GOV)

Snowberg, D.; Dana, S.; Hughes, S.

2014-09-01

A biaxial resonant test method was utilized to simultaneously fatigue test a wind turbine blade in the flap and edge (lead-lag) direction. Biaxial resonant blade fatigue testing is an accelerated life test method utilizing oscillating masses on the blade; each mass is independently oscillated at the respective flap and edge blade resonant frequency. The flap and edge resonant frequency were not controlled, nor were they constant for this demonstrated test method. This biaxial resonant test method presented surmountable challenges in test setup simulation, control and data processing. Biaxial resonant testing has the potential to complete test projects faster than single-axismore » testing. The load modulation during a biaxial resonant test may necessitate periodic load application above targets or higher applied test cycles.« less

Methods and systems to enhance flame holding in a gas turbine engine

DOEpatents

Zuo, Baifang [Simpsonville, SC; Lacy, Benjamin Paul [Greer, SC; Stevenson, Christian Xavier [Inman, SC

2012-01-31

A fuel nozzle including a swirler assembly that includes a shroud, a hub, and a plurality of vanes extending between the shroud and the hub. Each vane includes a pressure sidewall and an opposite suction sidewall coupled to the pressure sidewall at a leading edge and at a trailing edge. At least one suction side fuel injection orifice is formed adjacent to the leading edge and extends from a first fuel supply passage to the suction sidewall. A fuel injection angle is oriented with respect to the suction sidewall. The suction side fuel injection orifice is configured to discharge fuel outward from the suction sidewall. At least one pressure side fuel injection orifice extends from a second fuel supply passage to the pressure sidewall and is substantially parallel to the trailing edge. The pressure side fuel injection orifice is configured to discharge fuel tangentially from the trailing edge.

Broadband Noise Reduction of a Low-Speed Fan Noise Using Trailing Edge Blowing

NASA Technical Reports Server (NTRS)

Sutliff, Daniel L.

2005-01-01

An experimental proof-of-concept test was conducted to demonstrate reduction of rotor-stator interaction noise through the use of rotor-trailing edge blowing. The velocity deficit from the viscous wake of the rotor blades was reduced by injecting air into the wake from a continuous trailing edge slot. Hollow blades with interior guide vanes create flow channels through which externally supplied air flows from the blade root to the trailing edge. A previous paper documented the substantial tonal reductions of this Trailing Edge Rotor Blowing (TERB) fan. This report documents the broadband characteristics of TERB. The Active Noise Control Fan (ANCF), located at the NASA Glenn Research Center, was used as the proof-of-concept test bed. Two-component hotwire data behind the rotor, unsteady surface pressures on the stator vane, and farfield directivity acoustic data were acquired at blowing rates of 1.1, 1.5, and 1.8 percent of the total fan mass flow. The results indicate a substantial reduction in the rotor wake turbulent velocity and in the stator vane unsteady surface pressures. Based on the physics of the noise generation, these indirect measurements indicate the prospect of broadband noise reduction. However, since the broadband noise generated by the ANCF is rotor-dominated, any change in the rotor-stator interaction broadband noise levels is barely distinguishable in the farfield measurements.

Aerodynamic analysis of seamless horizontal stabilizer

NASA Astrophysics Data System (ADS)

Nithya, S.; Kanimozhi, S.

2017-05-01

This project presents an investigative view into the concept of seamless aeroelastic wing and hingeless flexible trailing edge. Wings are designed to provide maximum lift and minimal drag and weight. But with conventional wings where rivets are used and the control surfaces are separately hinged, parasite drag comes into play. This project is about analysing a smooth seamless wing with hinge-less flexible trailing edge. This type of wing reduces the drag considerably and the hinge-less trailing edge leads to a minimal control demand and reduces the noise produced when the aircraft comes for landing. Seamless aeroelastic wing will function as an integrated one piece lifting and control surface. It has been designed to enhance a desirable wing camber for control by deflecting a hinge-less flexible trailing edge part instead of a traditional hinged control surface. This kind of flexible wing can be achieved either by a curved beam and disc actuation mechanism or by piezo-electric materials, whose shape change can be achieved by electricity. The intent of this project is to analyze the effects of introducing the concept of Seamless Wing to the horizontal stabilizer. While the removal of rivets and serrations that hinge the elevators to the stabilizer reduces the overall drag by a reasonable value, the overall concept of a control surface-less stabilizer where the maneuvers are done by deflecting the trailing edge offers better maneuverability.

A three dimensional unsteady iterative panel method with vortex particle wakes and boundary layer model for bio-inspired multi-body wings

NASA Astrophysics Data System (ADS)

Dhruv, Akash; Blower, Christopher; Wickenheiser, Adam M.

2015-03-01

The ability of UAVs to operate in complex and hostile environments makes them useful in military and civil operations concerning surveillance and reconnaissance. However, limitations in size of UAVs and communication delays prohibit their operation close to the ground and in cluttered environments, which increase risks associated with turbulence and wind gusts that cause trajectory deviations and potential loss of the vehicle. In the last decade, scientists and engineers have turned towards bio-inspiration to solve these issues by developing innovative flow control methods that offer better stability, controllability, and maneuverability. This paper presents an aerodynamic load solver for bio-inspired wings that consist of an array of feather-like flaps installed across the upper and lower surfaces in both the chord- and span-wise directions, mimicking the feathers of an avian wing. Each flap has the ability to rotate into both the wing body and the inbound airflow, generating complex flap configurations unobtainable by traditional wings that offer improved aerodynamic stability against gusting flows and turbulence. The solver discussed is an unsteady three-dimensional iterative doublet panel method with vortex particle wakes. This panel method models the wake-body interactions between multiple flaps effectively without the need to define specific wake geometries, thereby eliminating the need to manually model the wake for each configuration. To incorporate viscous flow characteristics, an iterative boundary layer theory is employed, modeling laminar, transitional and turbulent regions over the wing's surfaces, in addition to flow separation and reattachment locations. This technique enables the boundary layer to influence the wake strength and geometry both within the wing and aft of the trailing edge. The results obtained from this solver are validated using experimental data from a low-speed suction wind tunnel operating at Reynolds Number 300,000. This method enables fast and accurate assessment of aerodynamic loads for initial design of complex wing configurations compared to other methods available.

Glove box shield

DOEpatents

Brackenbush, Larry W.; Hoenes, Glenn R.

1981-01-01

According to the present invention, a shield for a glove box housing radioactive material is comprised of spaced apart clamping members which maintain three overlapping flaps in place therebetween. There is a central flap and two side flaps, the side flaps overlapping at the interior edges thereof and the central flap extending past the intersection of the side flaps in order to insure that the shield is always closed when the user withdraws his hand from the glove box. Lead loaded neoprene rubber is the preferred material for the three flaps, the extent of lead loading depending upon the radiation levels within the glove box.

Active Management of Flap-Edge Trailing Vortices

NASA Technical Reports Server (NTRS)

Greenblatt, David; Yao, Chung-Sheng; Vey, Stefan; Paschereit, Oliver C.; Meyer, Robert

2008-01-01

The vortex hazard produced by large airliners and increasingly larger airliners entering service, combined with projected rapid increases in the demand for air transportation, is expected to act as a major impediment to increased air traffic capacity. Significant reduction in the vortex hazard is possible, however, by employing active vortex alleviation techniques that reduce the wake severity by dynamically modifying its vortex characteristics, providing that the techniques do not degrade performance or compromise safety and ride quality. With this as background, a series of experiments were performed, initially at NASA Langley Research Center and subsequently at the Berlin University of Technology in collaboration with the German Aerospace Center. The investigations demonstrated the basic mechanism for managing trailing vortices using retrofitted devices that are decoupled from conventional control surfaces. The basic premise for managing vortices advanced here is rooted in the erstwhile forgotten hypothesis of Albert Betz, as extended and verified ingeniously by Coleman duPont Donaldson and his collaborators. Using these devices, vortices may be perturbed at arbitrarily long wavelengths down to wavelengths less than a typical airliner wingspan and the oscillatory loads on the wings, and hence the vehicle, are small. Significant flexibility in the specific device has been demonstrated using local passive and active separation control as well as local circulation control via Gurney flaps. The method is now in a position to be tested in a wind tunnel with a longer test section on a scaled airliner configuration. Alternatively, the method can be tested directly in a towing tank, on a model aircraft, a light aircraft or a full-scale airliner. The authors believed that this method will have significant appeal from an industry perspective due to its retrofit potential with little to no impact on cruise (devices tucked away in the cove or retracted); low operating power requirements; small lift oscillations when deployed in a time-dependent manner; and significant flexibility with respect to the specific devices selected.

Generalization of analytical tools for helicopter-rotor airfoils

NASA Technical Reports Server (NTRS)

Gibbs, E. H.

1979-01-01

A state-of-the-art finite difference boundary-layer program incorporated into the NYU Transonic Analysis Program is described. Some possible treatments for the trailing edge region were investigated. Findings indicate the trailing edge region, still within the scope of an iterative potential flow, boundary layer program, appears feasible.

Cooling circuit for steam and air-cooled turbine nozzle stage

DOEpatents

Itzel, Gary Michael; Yu, Yufeng

2002-01-01

The turbine vane segment includes inner and outer walls with a vane extending therebetween. The vane includes leading and trailing edge cavities and intermediate cavities. An impingement plate is spaced from the outer wall to impingement-cool the outer wall. Post-impingement cooling air flows through holes in the outer wall to form a thin air-cooling film along the outer wall. Cooling air is supplied an insert sleeve with openings in the leading edge cavity for impingement-cooling the leading edge. Holes through the leading edge afford thin-film cooling about the leading edge. Cooling air is provided the trailing edge cavity and passes through holes in the side walls of the vane for thin-film cooling of the trailing edge. Steam flows through a pair of intermediate cavities for impingement-cooling of the side walls. Post-impingement steam flows to the inner wall for impingement-cooling of the inner wall and returns the post-impingement cooling steam through inserts in other intermediate cavities for impingement-cooling the side walls of the vane.

Study on measurement of leading and trailing edges of blades based on optical scanning system

NASA Astrophysics Data System (ADS)

Chao, Bi; Liu, Hongguang; Bao, Longxiang; Li, Di

2017-10-01

In the field of aeronautics, the geometry and dimensional accuracy of the blade edges has a large influence on the aerodynamic performance of aero engine. Therefore, a non-contact optical scanning system is established to realize the measurement of leading and trailing edges of blades in a rapid, precise and efficient manner in the paper. Based on the mechanical framework of a traditional CMM, the system is equipped with a specified sensing device as the scanning probe, which is made up by two new-style laser scanning sensors installed at a certain angle to each other by a holder. In the measuring procedure, the geometric dimensions of the measured blade edges on every contour plane are determined by the contour information on five transversals at the leading or trailing edges, which can be used to determine the machining allowance of the blades. In order to verify the effectiveness and practicality of the system set up, a precision forging blade after grinded is adopted as the measured object and its leading and trailing edges are measured by the system respectively. In the experiment, the thickness of blade edges on three contour planes is measured by the optical scanning system several times. As the experiment results show, the repeatability accuracy of the system can meet its design requirements and the inspecting demands of the blade edges. As a result, the optical scanning system could serve as a component of the intelligent manufacturing system of blades to improve the machining quality of the blade edges.

Decomposing the aerodynamic forces of low-Reynolds flapping airfoils

NASA Astrophysics Data System (ADS)

Moriche, Manuel; Garcia-Villalba, Manuel; Flores, Oscar

2016-11-01

We present direct numerical simulations of flow around flapping NACA0012 airfoils at relatively small Reynolds numbers, Re = 1000 . The simulations are carried out with TUCAN, an in-house code that solves the Navier-Stokes equations for an incompressible flow with an immersed boundary method to model the presence of the airfoil. The motion of the airfoil is composed of a vertical translation, heaving, and a rotation about the quarter of the chord, pitching. Both motions are prescribed by sinusoidal laws, with a reduced frequency of k = 1 . 41 , a pitching amplitude of 30deg and a heaving amplitude of one chord. Both, the mean pitch angle and the phase shift between pitching and heaving motions are varied, to build a database with 18 configurations. Four of these cases are analysed in detail using the force decomposition algorithm of Chang (1992) and Martín Alcántara et al. (2015). This method decomposes the total aerodynamic force into added-mass (translation and rotation of the airfoil), a volumetric contribution from the vorticity (circulatory effects) and a surface contribution proportional to viscosity. In particular we will focus on the second, analysing the contribution of the leading and trailing edge vortices that typically appear in these flows. This work has been supported by the Spanish MINECO under Grant TRA2013-41103-P. The authors thankfully acknowledge the computer resources provided by the Red Española de Supercomputacion.

Supersonic aerodynamic characteristics of a proposed Assured Crew Return Capability (ACRC) lifting-body configuration

NASA Technical Reports Server (NTRS)

Ware, George M.

1989-01-01

An investigation was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers from 1.6 to 4.5. The model had a low-aspect-ratio body with a flat undersurface. A center fin and two outboard fins were mounted on the aft portion of the upper body. The outboard fins were rolled outboard 40 deg from the vertical. Elevon surfaces made up the trailing edges of the outboard fins, and body flaps were located on the upper and lower aft fuselage. The center fin pivoted about its midchord for yaw control. The model was longitudinally stable about the design center-of-gravity position at 54 percent of the body length. The configuration with undeflected longitudinal controls trimmed near 0 deg angle of attack at Mach numbers from 1.6 to 3.0 where lift and lift-drag ratio were negative. Longitudinal trim was near the maximum lift-drag ratio (1.4) at Mach 4.5. The model was directionally stable over Mach number range except at angles of attack around 4 deg at M = 2.5. Pitch control deflection of more than -10 deg with either elevons or body flaps is needed to trim the model to angles of attack at which lift becomes positive. With increased control deflection, the lifting-body configuration should perform the assured crew return mission through the supersonic speed range.

Advanced natural laminar flow airfoil with high lift to drag ratio

NASA Technical Reports Server (NTRS)

Viken, Jeffrey K.; Pfenninger, Werner; Mcghee, Robert J.

1986-01-01

An experimental verification of a high performance natural laminar flow (NLF) airfoil for low speed and high Reynolds number applications was completed in the Langley Low Turbulence Pressure Tunnel (LTPT). Theoretical development allowed for the achievement of 0.70 chord laminar flow on both surfaces by the use of accelerated flow as long as tunnel turbulence did not cause upstream movement of transition with increasing chord Reynolds number. With such a rearward pressure recovery, a concave type deceleration was implemented. Two-dimensional theoretical analysis indicated that a minimum profile drag coefficient of 0.0026 was possible with the desired laminar flow at the design condition. With the three-foot chord two-dimensional model constructed for the LTPT experiment, a minimum profile drag coefficient of 0.0027 was measured at c sub l = 0.41 and Re sub c = 10 x 10 to the 6th power. The low drag bucket was shifted over a considerably large c sub l range by the use of the 12.5 percent chord trailing edge flap. A two-dimensional lift to drag ratio (L/D) was 245. Surprisingly high c sub l max values were obtained for an airfoil of this type. A 0.20 chort split flap with 60 deg deflection was also implemented to verify the airfoil's lift capabilities. A maximum lift coefficient of 2.70 was attained at Reynolds numbers of 3 and 6 million.

Thrust Removal Scheme for the FAST-MAC Circulation Control Model Tested in the National Transonic Facility

NASA Technical Reports Server (NTRS)

Chan, David T.; Milholen, William E., II; Jones, Gregory S.; Goodliff, Scott L.

2014-01-01

A second wind tunnel test of the FAST-MAC circulation control semi-span model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model allowed independent control of four circulation control plenums producing a high momentum jet from a blowing slot near the wing trailing edge that was directed over a 15% chord simple-hinged flap. The model was configured for transonic testing of the cruise configuration with 0deg flap deflection to determine the potential for drag reduction with the circulation control blowing. Encouraging results from analysis of wing surface pressures suggested that the circulation control blowing was effective in reducing the transonic drag on the configuration, however this could not be quantified until the thrust generated by the blowing slot was correctly removed from the force and moment balance data. This paper will present the thrust removal methodology used for the FAST-MAC circulation control model and describe the experimental measurements and techniques used to develop the methodology. A discussion on the impact to the force and moment data as a result of removing the thrust from the blowing slot will also be presented for the cruise configuration, where at some Mach and Reynolds number conditions, the thrust-removed corrected data showed that a drag reduction was realized as a consequence of the blowing.

Design & fabrication of two seated aircraft with an advanced rotating leading edge wing

NASA Astrophysics Data System (ADS)

Al Ahmari, Saeed Abdullah Saeed

The title of this thesis is "Design & Fabrication of two Seated Aircraft with an Advanced Rotating Leading Edge Wing", this gives almost a good description of the work has been done. In this research, the moving surface boundary-layer control (MSBC) concept was investigated and implemented. An experimental model was constructed and tested in wind tunnel to determine the aerodynamic characteristics using the leading edge moving surface of modified semi-symmetric airfoil NACA1214. The moving surface is provided by a high speed rotating cylinder, which replaces the leading edge of the airfoil. The angle of attack, the cylinder surfaces velocity ratio Uc/U, and the flap deflection angle effects on the lift and drag coefficients and the stall angle of attack were investigated. This new technology was applied to a 2-seat light-sport aircraft that is designed and built in the Aerospace Engineering Department at KFUPM. The project team is led by the aerospace department chairman Dr. Ahmed Z. AL-Garni and Dr. Wael G. Abdelrahman and includes graduate and under graduate student. The wing was modified to include a rotating cylinder along the leading edge of the flap portion. This produced very promising results such as the increase of the maximum lift coefficient at Uc/U=3 by 82% when flaps up and 111% when flaps down at 40° and stall was delayed by 8degrees in both cases. The laboratory results also showed that the effective range of the leading-edge rotating cylinder is at low angles of attack which reduce the need for higher angles of attack for STOL aircraft.

Turbine blade squealer tip rail with fence members

DOE Office of Scientific and Technical Information (OSTI.GOV)

Little, David A

2012-11-20

A turbine blade includes an airfoil, a blade tip section, a squealer tip rail, and a plurality of chordally spaced fence members. The blade tip section includes a blade tip floor located at an end of the airfoil distal from the root. The blade tip floor includes a pressure side and a suction side joined together at chordally spaced apart leading and trailing edges of the airfoil. The squealer tip rail extends radially outwardly from the blade tip floor adjacent to the suction side and extends from a first location adjacent to the airfoil trailing edge to a second locationmore » adjacent to the airfoil leading edge. The fence members are located between the airfoil leading and trailing edges and extend radially outwardly from the blade tip floor and axially from the squealer tip rail toward the pressure side.« less

Off-Design Performance of Radial-Inflow Turbines

NASA Technical Reports Server (NTRS)

Meitner, P. L.; Glassman, A. J.

1986-01-01

Computer code determines rotor exit flow from hub to tip. RTOD (Radial Turbine Off-Design), computes off-design performance of radial turbine by modeling flow with stator viscous and trailing-edge losses, and with vaneless space loss between stator and rotor, and with rotor incidence, viscous, clearance, trailing-edge, and disk friction losses.

Optimal perturbations of a finite-width mixing layer near the trailing edge

NASA Astrophysics Data System (ADS)

Gumbart, James C.; Rabchuk, James

2002-03-01

The trailing edge of a surface separating two fluid flows can act as an efficient receptor for acoustic or other disturbances. The incident wave energy is converted by a linear mechanism into incipient flow instabilities which lead further downstream to the transition to turbulence. Understanding this process is essential for analyzing feedback loops and other resonances which can cause unwanted structural vibrations in the surface material or directed acoustic emissions from the mixing region. Previously, the modes of instability in a finite-width mixing layer near the trailing edge were studied as a function of frequency by assuming that vorticity was continually being introduced into the flow at the trailing edge by the forcing field. It was found that the initial amplitude of the growing instability mode was a sharply decreasing function of forcing frequency, and that the initial amplitude was a minimum for the frequency at which the rate of instability growth was a maximum^1. This result has led to a study of the adjoint equation for the perturbation stream function, whose eigensolutions are known to be associated with the optimal perturbation field for the frequency of forcing leading to the greatest instability growth downstream. We have obtained these solutions for a piecewise linear velocity profile near the trailing edge using group-theoretic techniques and have shown that they are indeed optimal. We have also analyzed the nature of the physical forcing field that might produce these optimal perturbations. ^1 Rabchuk, J.A., July 2000, Physics of Fluids.

Receptivity to Unsteady Disturbances at the Trailing Edge in a Finite-Width Mixing Layer Flow.

NASA Astrophysics Data System (ADS)

Rabchuk, James Allen

1995-01-01

A theoretical study of the receptivity to harmonic disturbances at the trailing edge of a finite-width mixing layer has been carried out. The unsteady Kutta condition at the trailing edge has been reexamined at the vorticity scale of the steady mixing layer profile, and the underlying physical mechanism of this condition explained. The receptivity problem of harmonic forcing at the trailing edge is shown to reduce to an initial-value problem for the downstream mixing layer or wake. A linear coupling term for the response field amplitude is derived which is proportional to the square root of the Strouhal number and the difference in the gradient of the forcing pressure field tangential to the plate near the trailing edge. An initial-value problem is then solved for an inviscid, incompressible mixing layer with a piecewise linear velocity profile leaving the trailing edge of a flat plate, subject to harmonic forcing. The Wiener-Hopf technique is used to solve for the stream function of the response field over a range of forcing frequencies and mean flow velocities. The solutions are shown to agree with previous solutions for infinitesimally thin shear layers from Bechert, 1988 and Orszag and Crow, 1970, in the limit that the Strouhal number relative to the mixing layer thickness, S, is small. In addition, solutions are obtained for moderate values of S, for which the mixing layer is most unstable. It is shown that for increasing S, the initial amplitudes of the discrete modes of instability decrease like 1 over S and then level off, while the neutrally stable mode of response is increasingly amplified. It is also shown that the overall phase of the response is nearly independent of S, except at a cross-stream position where the phase shifts by 180 degrees and the amplitude of the response goes to zero, which moves from the low to the high speed flow as S increases.

Lift producing device exhibiting low drag and reduced ventilation potential and method for producing the same

NASA Technical Reports Server (NTRS)

Caldwell, Richard A. (Inventor)

1991-01-01

A lift producing device is disclosed which is adapted to be connected to a vehicle to provide lift to the vehicle when the vehicle is moved relative to a first fluid medium having a first density and viscosity and being in contact with a second fluid medium adjacent the vehicle. The second fluid medium has a second fluid density which is different from the first fluid density. The lift producing device comprises opposed first and second major surfaces joined at a longitudinally extending leading edge and at a longitudinally extending trailing edge, with at least a portion of the longitudinally extending leading edge being spaced from the longitudinally extending trailing edge by a predetermined mean chord length. When the vehicle is moved relative to the first fluid medium at a velocity within a range of predetermined velocities, with each of the velocities having a direction inclined from a plane extending through the leading edge and the trailing edge within a predetermined angular range, a region of high pressure is generated in the first fluid medium adjacent the first major surface and a region of low pressure is generated in the first fluid medium adjacent the second major surface. The lift producing device has a cross-sectional shape which will generate a pressure distribution around the device when the vehicle is moved relative to the first fluid medium at a velocity within the range of predetermined velocities such that the first fluid medium exhibits attached laminar flow along the device for a portion of the predetermined mean chord length from the leading edge to the trailing edge and will neither form a laminar separation bubble adjacent the second major surface of the device, nor exhibit turbulent separation adjacent the second major surface for substantially all of the predetermined mean chord length from the leading edge to the trailing edge. The portion along which attached laminar flow is maintained is the longest portion which will still fulfill the flow separation requirements. A method for producing the foil is also disclosed.

Effect of Trailing Edge Flow Injection on Fan Noise and Aerodynamic Performance

NASA Technical Reports Server (NTRS)

Fite, E. Brian; Woodward, Richard P.; Podboy, Gary G.

2006-01-01

An experimental investigation using trailing edge blowing for reducing fan rotor/guide vane wake interaction noise was completed in the NASA Glenn 9- by 15-foot Low Speed Wind Tunnel. Data were acquired to measure noise, aerodynamic performance, and flow features for a 22" tip diameter fan representative of modern turbofan technology. The fan was designed to use trailing edge blowing to reduce the fan blade wake momentum deficit. The test objective was to quantify noise reductions, measure impacts on fan aerodynamic performance, and document the flow field using hot-film anemometry. Measurements concentrated on approach, cutback, and takeoff rotational speeds as those are the primary conditions of acoustic interest. Data are presented for a 2% (relative to overall fan flow) trailing edge injection rate and show a 2 dB reduction in Overall Sound Power Level (OAPWL) at all fan test speeds. The reduction in broadband noise is nearly constant and is approximately 1.5 dB up to 20 kHz at all fan speeds. Measurements of tone noise show significant variation, as evidenced by reductions of up to 6 dB in the 2 BPF tone at 6700 rpm.: and increases of nearly 2 dB for the 4 BPF tone at approach speed. Aerodynamic performance measurements show the fan with 2 % injection has an overall efficiency that is comparable to the baseline fan and operates, as intended, with nearly the same pressure ratio and mass flow parameters. Hot-film measurements obtained at the approach operating condition indicate that mean blade wake filling in the tip region was not as significant as expected. This suggests that additional acoustic benefits could be realized if the trailing edge blowing could be modified to provide better filling of the wake momentum deficit. Nevertheless, the hot-film measurements indicate that the trailing edge blowing provided significant reductions in blade wake turbulence. Overall, these results indicate that further work may be required to fully understand the proper implementation of injecting flow at/near the trailing edge as a wake filling strategy. However, data do support the notion that noise reductions can be realized not only for tones but perhaps more importantly, also for broadband. Furthermore, the technique can be implemented without adversely effecting overall fan aerodynamic performance.

Parametric Evaluation of Thin, Transonic Circulation-Control Airfoils

NASA Technical Reports Server (NTRS)

Schlecht, Robin; Anders, Scott

2007-01-01

Wind-tunnel tests were conducted in the NASA Langley Transonic Dynamics Tunnel on a 6 percent-thick, elliptical circulation-control airfoil with upper-surface and lower-surface blowing capability. Results for elliptical Coanda trailing-edge geometries, biconvex Coanda trailing-edge geometries, and leading-edge geometries are reported. Results are presented at subsonic and transonic Mach numbers of 0.3 and 0.8, respectively. When considering one fixed trailing-edge geometry, for both the subsonic and transonic conditions it was found that the [3.0:1] ratio elliptical Coanda surface with the most rounded leading-edge [03] performed favorably and was determined to be the best compromise between comparable configurations that took advantage of the Coanda effect. This configuration generated a maximum. (Delta)C(sub 1) = 0.625 at a C(sub mu) = 0.06 at M = 0.3, alpha = 6deg. This same configuration generated a maximum (Delta)C(sub 1) = 0.275 at a C(sub mu) = 0.0085 at M = 0.8, alpha = 3deg.

[Low-Frequency Flow Oscillation

NASA Technical Reports Server (NTRS)

Bragg, Michael B.

1997-01-01

The results of the research conducted under this grant are presented in detail in three Master theses, by Heinrich, Balow, and Broeren. Additional analysis of the experimental data can be found in two AIAA Journal articles and two conference papers. Citations for all of the studies' publications can be found in the bibliography which is attached. The objective of Heinrich's study was to document the low-frequency flow oscillation on the LRN-1007 airfoil, which had been previously observed at low Reynolds number, to determine its origin, and explore the phenomenon at higher Reynolds number. Heinrich performed detailed flow visualization on the airfoil using surface fluorescent oil and laser-sheet off-body visualization. A large leading-edge separation bubble and trailing-edge separation was identified on the airfoil just prior to the onset of the unsteady stall flow oscillation. From the laser-sheet data, the unsteady flow appeared as a massive boundary-layer separation followed by flow reattachment. Hot-wire data were taken in the wake to identify the presence of the flow oscillation and the dominant frequency. The oscillation was found in the flow from a Reynolds number of 0.3 to 1.3 x 10 exp 6. The Strouhal number based on airfoil projected height was nominally 0.02 and increased slightly with increasing Reynolds number and significantly with increasing airfoil angle of attack. Balow focused his research on the leading-edge separation bubble which was hypothesized to be the origin of the low-frequency oscillation. Initially, experimental measurements in the bubble at the onset of the low-frequency oscillation were attempted to study the characteristics of the bubble and explain possible relationships to the shear-layer-flapping phenomena. Unfortunately, the bubble proved to be extremely sensitive to the probe interference and it drastically reduced the size of the bubble. These detailed measurements were then abandoned by Balow. However, this led to a series of tests where the leading-edge bubble and trailing-edge separation were altered and the affect on the flow-oscillation studied. Balow found that by tripping the airfoil boundary-layer with "zigzag" tape ahead of bubble separation, the bubble was effectively eliminated mid the oscillation suppressed. Wake survey drag measurements showed a drastic reduction in airfoil drag when the bubble and oscillation were eliminated. Using the "zigzag" tape, the trailing-edge separation was moved downstream approximately 5 percent chord. This was found to reduce the amplitude of the oscillation, particularly in the onset stage at low angle of attack (around 14 degrees). Through detailed analysis of the wake behind the airfoil during the unsteady flow oscillation, Balow provided a better understanding of the wake flowfield. Broeren studied the oscillating flowfield in detail at Reynolds number equal 3 x 10 exp 5 and an angle of attack of 15 degrees using laser Doppler velocimetry (LDV). Two-dimensional LDV data were acquired at 687 grid points above the model upper surface while hot-wire data were taken simultaneously in the wake. Using the hot-wire signal, the LDV data were phase averaged into 24 bins to represent a single ensemble average of one oscillation cycle. The velocity data showed a flowfield oscillation that could be divided into three flow regimes. In the first regime, the flow over the airfoil was completely separated initially, the flowfield reattached from the leading edge and the reattachment point moved downstream with increasing time or phase. Broeren referred to this as the reattachment regime. The bubble development regime followed, where a leading-edge separation bubble formed at the leading edge and grew with increasing time. During the initial part of this regime the trailing-edge separation continued to move downstream. However, during the last 30 degrees of phase the trailing-edge separation moved rapidly forward and appeared to merge with the leading-edge bubble. During the third regime, the separation regime, the flow was segmented from the airfoil leading edge and did not reattach to the airfoil surface. The reverse flow was seen to grow in vertical extent up from the model surface as the phase increased. Next reattachment began again at the leading edge signaling the start of the reattachment regime, and so the cycle continued. From Broeren's work, the details of the unsteady flowfield over the airfoil were seen for the first time. From this research a great deal has been learned about the low-frequency flow oscillation which naturally occurs on the LRN-1007 airfoil near stall. The oscillation was seen to persist at higher Reynolds number, the dependence of the Strouhal number on angle of attack and Reynolds number were discovered, the critical role played by the laminar bubble was shown and the entire upper surface flowfield during a flow oscillation cycle was measured and analyzed. What still eludes understanding is the scaling of the flow oscillation and why certain airfoils, such as the LRN, have a very strong low-frequency mode and other airfoils exhibit no organized low-frequency oscillation at all.

Crescentic ramp turbine stage

NASA Technical Reports Server (NTRS)

Lee, Ching-Pang (Inventor); Tam, Anna (Inventor); Kirtley, Kevin Richard (Inventor); Lamson, Scott Henry (Inventor)

2007-01-01

A turbine stage includes a row of airfoils joined to corresponding platforms to define flow passages therebetween. Each airfoil includes opposite pressure and suction sides and extends in chord between opposite leading and trailing edges. Each platform includes a crescentic ramp increasing in height from the leading and trailing edges toward the midchord of the airfoil along the pressure side thereof.

An experimental study of a turbulent boundary layer in the trailing edge region of a circulation-control airfoil

NASA Technical Reports Server (NTRS)

Heinemann, K.; Brown, Jeff

1992-01-01

This report discusses progress made on NASA Cooperative Agreement NCC2-545, 'An Experimental Study of a Turbulent Boundary Layer in the Trailing-Edge Region of a Circulation-Control Airfoil' during the period 9/1/91 through 9/30/92. The study features 2-component laser Doppler velocimeter (LDV) measurements in the trailing edge and wake regions of a generic 2-dimensional circulation-control model. The final experimental phase of the study will be carried out in the Ames High Reynolds Number Channel 2 (HRC2) transonic blow-down-facility. During the 13-month period covered by this report, work continued on the development of the near-wall laser Doppler velocimeter (LDV) described in previous reports.

Downstream Effects on Orbiter Leeside Flow Separation for Hypersonic Flows

NASA Technical Reports Server (NTRS)

Buck, Gregory M.; Pulsonetti, Maria V.; Weilmuenster, K. James

2005-01-01

Discrepancies between experiment and computation for shuttle leeside flow separation, which came to light in the Columbia accident investigation, are resolved. Tests were run in the Langley Research Center 20-Inch Hypersonic CF4 Tunnel with a baseline orbiter model and two extended trailing edge models. The extended trailing edges altered the wing leeside separation lines, moving the lines toward the fuselage, proving that wing trailing edge modeling does affect the orbiter leeside flow. Computations were then made with a wake grid. These calculations more closely matched baseline experiments. Thus, the present findings demonstrate that it is imperative to include the wake flow domain in CFD calculations in order to accurately predict leeside flow separation for hypersonic vehicles at high angles of attack.

Application of Excitation from Multiple Locations on a Simplified High-Lift System

NASA Technical Reports Server (NTRS)

Melton, LaTunia Pack; Yao, Chung-Sheng; Seifert, Avi

2004-01-01

A series of active flow control experiments were recently conducted on a simplified high-lift system. The purpose of the experiments was to explore the prospects of eliminating all but simply hinged leading and trailing edge flaps, while controlling separation on the supercritical airfoil using multiple periodic excitation slots. Excitation was provided by three. independently controlled, self-contained, piezoelectric actuators. Low frequency excitation was generated through amplitude modulation of the high frequency carrier wave, the actuators' resonant frequencies. It was demonstrated, for the first time, that pulsed modulated signal from two neighboring slots interact favorably to increase lift. Phase sensitivity at the low frequency was measured, even though the excitation was synthesized from the high-frequency carrier wave. The measurements were performed at low Reynolds numbers and included mean and unsteady surface pressures, surface hot-films, wake pressures and particle image velocimetry. A modest (6%) increase in maximum lift (compared to the optimal baseline) was obtained due t o the activation of two of the three actuators.

Aerodynamic Modeling of Transonic Aircraft Using Vortex Lattice Coupled with Transonic Small Disturbance for Conceptual Design

NASA Technical Reports Server (NTRS)

Chaparro, Daniel; Fujiwara, Gustavo E. C.; Ting, Eric; Nguyen, Nhan

2016-01-01

The need to rapidly scan large design spaces during conceptual design calls for computationally inexpensive tools such as the vortex lattice method (VLM). Although some VLM tools, such as Vorview have been extended to model fully-supersonic flow, VLM solutions are typically limited to inviscid, subcritical flow regimes. Many transport aircraft operate at transonic speeds, which limits the applicability of VLM for such applications. This paper presents a novel approach to correct three-dimensional VLM through coupling of two-dimensional transonic small disturbance (TSD) solutions along the span of an aircraft wing in order to accurately predict transonic aerodynamic loading and wave drag for transport aircraft. The approach is extended to predict flow separation and capture the attenuation of aerodynamic forces due to boundary layer viscosity by coupling the TSD solver with an integral boundary layer (IBL) model. The modeling framework is applied to the NASA General Transport Model (GTM) integrated with a novel control surface known as the Variable Camber Continuous Trailing Edge Flap (VCCTEF).

Analysis of a high speed civil transport configuration at subsonic flow conditions using a Navier-Stokes solver

NASA Technical Reports Server (NTRS)

Lessard, Victor R.

1993-01-01

Computations of three dimensional vortical flows over a generic High Speed Civil Transport (HSCT) configuration with an aspect ratio of 3.04 are performed using a thin-layer Navier-Stokes solver. The HSCT cruise configuration is modeled without leading or trailing edge flap deflections and without engine nacelles. The flow conditions, which correspond to tests done in the NASA Langley 8-Foot Transonic Pressure Tunnel (TPT), are a subsonic Mach number of 0.3 and Reynolds number of 4.4 million for a range-of-attack (-.23 deg to 17.78 deg). The effects of the farfield boundary location with respect to the body are investigated. The boundary layer is assumed turbulent and simulated using an algebraic turbulence model. The key features of the vortices and their interactions are captured. Grid distribution in the vortex regions is critical for predicting the correct induced lift. Computed forces and surface pressures compare reasonably well with the experimental TPT data.

Flow speed has little impact on propulsive characteristics of oscillating foils

NASA Astrophysics Data System (ADS)

Van Buren, T.; Floryan, D.; Wei, N.; Smits, A. J.

2018-01-01

Experiments are reported on the performance of a pitching and heaving two-dimensional foil in a water channel in either continuous or intermittent motion. We find that the thrust and power are independent of the mean free-stream velocity for twofold changes in the mean velocity (fourfold in the dynamic pressure) and for oscillations in the velocity up to 38% of the mean, where the oscillations are intended to mimic those of freely swimming motions where the thrust varies during the flapping cycle. We demonstrate that the correct velocity scale is not the flow velocity but the mean velocity of the trailing edge. We also find little or no impact of streamwise velocity change on the wake characteristics such as vortex organization, vortex strength, and time-averaged velocity profile development—the wake is both qualitatively and quantitatively unchanged. Our results suggest that constant velocity studies can be used to make robust conclusions about swimming performance without a need to explore the free-swimming condition.