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At least 1,387 records · Page 77
All sourcesNASA NTRSDOE OSTI
NASA NTRS1988

A static investigation of a simultaneous pitch and yaw thrust vectoring 2-D C-D nozzle

An investigation has been conducted in the static test facility of the Langley 16-Foot Transonic Tunnel to determine the internal performance and flow-turning capability of a two-dimensional convergent-divergent nozzle. Thrust vectoring in the pitch plane was provided by rotation of the divergent flaps. The exhaust stream was turned in the yaw direction by deflection of yaw flaps hinged at the end of the nozzle sidewalls. The yaw flap hinge location was varied along the divergent region of the nozzle at four locations including the exit plane and the throat plane. The three hinge locations upstream of the nozzle exit plane required the downstream corners of both upper and lower divergent flaps to be cut off to eliminate interference when the yaw flaps were deflected. Three different lengths of yaw flaps were tested at several angles of deflection. The nozzle simulated a dry power setting with an expansion ratio typical of subsonic cruise and was tested at nozzle pressure ratios from 2.0 to 7.0.

Taylor, John G.↗
NASA NTRS1991

Static performance of a multiaxis thrust vectoring cruciform nozzle

A multiaxis thrust vectoring nozzle designed to have equal flow turning capability in pitch and yaw is investigated, and methods for obtaining large multiaxis thrust vector angles without physical vectoring flap interference are studied. These methods include restricting the pitch flaps from the path of the yaw flaps, and shifting the flow path at the throat off the nozzle centerline in order to permit larger pitch flap deflections without interfering with operation of the yaw flaps. The results obtained show that unvectored performance of the cruciform nozzle is only slightly lower (0.5 to 1.0 percent) than that of previously tested axisymmetric and nonaxisymmetric nozzles, despite the complex internal geometry. The shifted-throat nozzle design has larger thrust vector angles at the nozzle pressure ratio of peak resultant thrust efficiency, but has 1 to 2 percent lower peak thrust performance than the restricted-flap nozzle design.

Wing, David J.↗
NASA NTRS1992

A mathematical model of a tilt-wing aircraft for piloted simulation

A mathematical model of a tilt-wing aircraft that was used in a piloted, six-degree-of-freedom flight simulation application is described. Two types of control systems developed for the math model are discussed: a conventional, programmed-flap wing-tilt control system and a geared-flap wing-tilt control system. The primary objective was to develop the capability to study tilt-wing aircraft. Experienced Tilt-wing pilots subjectively evaluated the model using programmed-flap control to assess the quality of the simulation. The math model was then applied to study geared-flap control to investigate the possibility of eliminating the need for auxilary pitch-control devices (such as the horizontal tail rotor or tail jet used in earlier tilt-wing designs). This investigation was performed in the moving-base simulation environment, and the vehicle responses with programmed-flap and geared-flap control were compared. The results of the evaluation of the math model are discussed.

Totah, Joseph J.↗
NASA NTRS1991

Description of a tilt wing mathematical model for piloted simulation

A tilt-wing mathematical model that was used in a piloted six-deg-of-freedom flight simulation application is presented. Two types of control systems developed for the model - a conventional programmed-flap wing-tilt control system and a geared-flap wing-tilt control system - are discussed. The objective of this effort was to develop the capability to study tilt-wing aircraft. Experienced tilt-wing pilots subjectively evaluated the model using programmed-flap control to assess the quality of the simulation. The objective was met and the model was then applied to study geared-flap control to investigate the possibility of eliminating the need for auxiliary pitch control devices. This was performed in the moving-base simulation environment, and the vehicle responses with programmed-flap and geared-flap control were compared.

Totah, Joseph J.↗
NASA NTRS1993

Static internal performance of a single expansion ramp nozzle with multiaxis thrust vectoring capability

An investigation was conducted at static conditions in order to determine the internal performance characteristics of a multiaxis thrust vectoring single expansion ramp nozzle. Yaw vectoring was achieved by deflecting yaw flaps in the nozzle sidewall into the nozzle exhaust flow. In order to eliminate any physical interference between the variable angle yaw flap deflected into the exhaust flow and the nozzle upper ramp and lower flap which were deflected for pitch vectoring, the downstream corners of both the nozzle ramp and lower flap were cut off to allow for up to 30 deg of yaw vectoring. The effects of nozzle upper ramp and lower flap cutout, yaw flap hinge line location and hinge inclination angle, sidewall containment, geometric pitch vector angle, and geometric yaw vector angle were studied. This investigation was conducted in the static-test facility of the Langley 16-Foot Transonic Tunnel at nozzle pressure ratios up to 8.0.

Capone, Francis J.↗
NASA NTRS1994

Dynamic ground effects flight test of an F-15 aircraft

Flight tests to determine the changes in the aerodynamic characteristics of an F-15 aircraft caused by dynamic ground effects are described. Data were obtained for low and high sink rates between 0.7 and 6.5 ft/sec and at two landing approach speeds and flap settings: 150 kn with the flaps down and 170 kn with the flaps up. Simple correlation curves are given for the change in aerodynamic coefficients because of ground effects as a function of sink rate. Ground effects generally caused an increase in the lift, drag, and nose-down pitching movement coefficients. The change in the lift coefficient increased from approximately 0.05 at the high-sink rate to approximately 0.10 at the low-sink rate. The change in the drag coefficient increased from approximately 0 to 0.03 over this decreasing sink rate range. No significant difference because of the approach configuration was evident for lift and drag; however, a significant difference in pitching movement was observed for the two approach speeds and flap settings. For the 170 kn with the flaps up configuration, the change in the nose-down pitching movement increased from approximately -0.008 to -0.016. For the 150 kn with the flaps down configuration, the change was approximately -0.008 to -0.038.

Corda, Stephen↗
NASA NTRS1995

Application of a Navier-Stokes Solver to the Analysis of Multielement Airfoils and Wings Using Multizonal Grid Techniques

A computational study was performed to determine the predictive capability of a Reynolds averaged Navier-Stokes code (CFL3D) for two-dimensional and three-dimensional multielement high-lift systems. Three configurations were analyzed: a three-element airfoil, a wing with a full span flap and a wing with a partial span flap. In order to accurately model these complex geometries, two different multizonal structured grid techniques were employed. For the airfoil and full span wing configurations, a chimera or overset grid technique was used. The results of the airfoil analysis illustrated that although the absolute values of lift were somewhat in error, the code was able to predict reasonably well the variation with Reynolds number and flap position. The full span flap analysis demonstrated good agreement with experimental surface pressure data over the wing and flap. Multiblock patched grids were used to model the partial span flap wing. A modification to an existing patched- grid algorithm was required to analyze the configuration as modeled. Comparisons with experimental data were very good, indicating the applicability of the patched-grid technique to analyses of these complex geometries.

Jones, Kenneth M.↗
NASA NTRS1994

Maneuver Performance Enhancement for an Advanced Fighter/Attack Aircraft

A small scale wind tunnel test of a realistic fighter configuration has been completed in NASA Ames' 7'x10' wind tunnel. This test was part of the Fighter Lift and Control (FLAC) program, a joint NASA - USAF research program, involving small and large-scale wind-tunnel tests and computational analysis of unique lift augmentation and control devices. The goal of this program is to enhance the maneuver and control capability of next-generation Air Force multi-role fighter aircraft with low-observables geometries. The principal objective of this test was to determine the effectiveness of passive boundary layer control devices at increasing L/D at sustained maneuver lift coefficients. Vortex generators (VGs) were used to energize the boundary layer to prevent or delay separation. Corotating vanes, counter-rotating vanes, and Wheeler Wishbone VGs were used in the vicinity of the leading and trailing edge flap hinge lines. Principle test parameters were leading and trailing edge flap deflections, and location, size, spacing, and orientation for each VG type. Gurney flaps were also tested. Data gathered include balance force and moment data, surface pressures, and flow visualization for characterizing flow behavior and locating separation lines. Results were quite different for the two best flap configurations tested. All VG types tested showed improvement (up to 5%) in maneuver L/D with flaps at LE=20 degrees, TE=0 degrees. The same VGs degraded performance, in all but a few cases, with flaps at LE=15 degrees, TE=10 degrees.

Samuels, Jeff↗
NASA NTRS1941

The Development of a New Lateral-Control Arrangement

Development work on an arrangement using ailerons and spoilers for lateral control was carried out by the Vought-Sikorsky Aircraft Division of the United Aircraft Corporation on a small commercial airplane in flight and on an airfoil in a wind tunnel. Spoiler hinge moments were reduced by aerodynamic balance. The arrangement was then built into an experimental airplane and further improvements were adopted as the result of flight and tunnel tests. The use of ailerons for lateral control with flaps up, spoilers with flaps full down, and gradual transition as the flaps are lowered was found to provide lateral control under the flight conditions for which they were best suited. The ailerons were of short span, permitting the use of long-span flaps, and were drooped to a relatively large angle when the flaps were deflected. A high maximum lift coefficient was thus attained. With large control deflections in the intermediate flap-angle range and spoiler effectiveness near neutral improved by "ventilating" the spoiler, the lateral control was satisfactory for the experimental airplane and was a definite improvement over that of a conventional control arrangement.

Baker, Paul S.↗
NASA NTRS1940

Comparison of Three Exit-Area Control Devices on an N.A.C.A. Cowling, Special Report

Adjustable cowling flaps, an adjustable-length cowling skirt, and a bottom opening with adjustable flap were tested as means of controlling the rate of cooling-air flow through an air-cooled radial-engine cowling. The devices were tested in the NACA 20-foot tunnel on a model wing-nacelle-propeller combination, through an airspeed range of 20 to 80 miles per hour, and with the propeller blade angle set 23 degrees at 0.75 of the tip radius. The resistance of the engine to air flow through the cowling was simulated by a perforated plate. The results indicated that the adjustable cowling flap and the bottom opening with adjustable flap were about equally effective on the basis of pressure drop obtainable and that both were more effective means of increasing the pressure drop through the cowling than the adjustable-length skirt. At conditions of equal cooling-air flow, the net efficiency obtained with the adjustable cowling flaps and the adjustable-length cowling skirt was about 1% greater than the net efficiency obtained with the bottom opening with adjustable flap.

McHugh, James G.↗
NASA NTRS2010

Simple Check Valves for Microfluidic Devices

A simple design concept for check valves has been adopted for microfluidic devices that consist mostly of (1) deformable fluorocarbon polymer membranes sandwiched between (2) borosilicate float glass wafers into which channels, valve seats, and holes have been etched. The first microfluidic devices in which these check valves are intended to be used are micro-capillary electrophoresis (microCE) devices undergoing development for use on Mars in detecting compounds indicative of life. In this application, it will be necessary to store some liquid samples in reservoirs in the devices for subsequent laboratory analysis, and check valves are needed to prevent cross-contamination of the samples. The simple check-valve design concept is also applicable to other microfluidic devices and to fluidic devices in general. These check valves are simplified microscopic versions of conventional rubber- flap check valves that are parts of numerous industrial and consumer products. These check valves are fabricated, not as separate components, but as integral parts of microfluidic devices. A check valve according to this concept consists of suitably shaped portions of a deformable membrane and the two glass wafers between which the membrane is sandwiched (see figure). The valve flap is formed by making an approximately semicircular cut in the membrane. The flap is centered over a hole in the lower glass wafer, through which hole the liquid in question is intended to flow upward into a wider hole, channel, or reservoir in the upper glass wafer. The radius of the cut exceeds the radius of the hole by an amount large enough to prevent settling of the flap into the hole. As in a conventional rubber-flap check valve, back pressure in the liquid pushes the flap against the valve seat (in this case, the valve seat is the adjacent surface of the lower glass wafer), thereby forming a seal that prevents backflow.

Willis, Peter A.↗
NASA NTRS2012

Aeroacoustic Study of a High-Fidelity Aircraft Model: Unsteady Surface Pressures - Part 2

In this paper, we present unsteady surface pressure measurements for an 18%-scale, semi-span Gulfstream aircraft model. This high-fidelity model is being used to perform detailed studies of airframe noise associated with main landing gear, flap components, and gear-flap interaction noise, as well as to evaluate novel noise reduction concepts. The aerodynamic segment of the tests, conducted in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel, was completed in November 2010. To discern the characteristics of the surface pressure fluctuations in the vicinity of the prominent noise sources, unsteady sensors were installed on the inboard and outboard flap edges, and on the main gear wheels, struts, and door. Various configurations were tested, including flap deflections of 0?, 20?, and 39?, with and without the main landing gear. The majority of unsteady surface pressure measurements were acquired for the nominal landing configuration where the main gear was deployed and the flap was deflected 39?. To assess the Mach number variation of the surface pressure amplitudes, measurements were obtained at Mach numbers of 0.16, 0.20, and 0.24. Comparison of the unsteady surface pressures with the main gear on and off shows significant interaction between the gear wake and the inboard flap edge, resulting in higher amplitude fluctuations when the gear is present.

Khorrami, Mehdi R.↗
NASA NTRS2014

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

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.

Melton, Latunia P.↗
NASA NTRS2015

Semi-Empirical Prediction of Aircraft Low-Speed Aerodynamic Characteristics

This paper lays out a comprehensive methodology for computing a low-speed, high-lift polar, without requiring additional details about the aircraft design beyond what is typically available at the conceptual design stage. Introducing low-order, physics-based aerodynamic analyses allows the methodology to be more applicable to unconventional aircraft concepts than traditional, fully-empirical methods. The methodology uses empirical relationships for flap lift effectiveness, chord extension, drag-coefficient increment and maximum lift coefficient of various types of flap systems as a function of flap deflection, and combines these increments with the characteristics of the unflapped airfoils. Once the aerodynamic characteristics of the flapped sections are known, a vortex-lattice analysis calculates the three-dimensional lift, drag and moment coefficients of the whole aircraft configuration. This paper details the results of two validation cases: a supercritical airfoil model with several types of flaps; and a 12-foot, full-span aircraft model with slats and double-slotted flaps.

Olson, Erik D.↗
NASA NTRS2016

Simulation-Based Airframe Noise Prediction of a Full-Scale, Full Aircraft

A previously validated computational approach applied to an 18%-scale, semi-span Gulfstream aircraft model was extended to the full-scale, full-span aircraft in the present investigation. The full-scale flap and main landing gear geometries used in the simulations are nearly identical to those flown on the actual aircraft. The lattice Boltzmann solver PowerFLOW® was used to perform time-accurate predictions of the flow field associated with this aircraft. The simulations were performed at a Mach number of 0.2 with the flap deflected 39 deg. and main landing gear deployed (landing configuration). Special attention was paid to the accurate prediction of major sources of flap tip and main landing gear noise. Computed farfield noise spectra for three selected baseline configurations (flap deflected 39 deg. with and without main gear extended, and flap deflected 0 deg. with gear deployed) are presented. The flap brackets are shown to be important contributors to the farfield noise spectra in the mid- to high-frequency range. Simulated farfield noise spectra for the baseline configurations, obtained using a Ffowcs Williams and Hawkings acoustic analogy approach, were found to be in close agreement with acoustic measurements acquired during the 2006 NASA-Gulfstream joint flight test of the same aircraft.

Khorrami, Mehdi R.↗
NASA NTRS2022

Computational Analysis of the X-57 Maxwell Airplane at Unpowered Conditions (Preliminary Fuselage)

The X-57 Maxwell is an all-electric airplane that implements a distributed electric propulsion system to demonstrate that high-efficiency electric propulsion can be integrated with aerodynamics to increase the performance of an airplane. To this end, distributed electric fans were installed on the wing to provide increased flow over the wing at the low takeoff and landing speeds of the X-57. The low-speed lift augmentation allows for a reduction in wing area for cruise optimization. The X-57 wing area was reduced to 42 percent of the wing area of the baseline aircraft, a Tecnam P2006T. With this reduced wing area and the electric propulsion system, it is estimated that the X-57 will cruise on less than one-third the total energy compared to the baseline aircraft. To meet the cruise performance goal at a Mach number of 0.233 at an altitude of 8000 feet, the X-57 has a cruise lift coefficient of 0.7516 and needs to have a cruise drag coefficient of 0.05423 or less. The USM3D computational solver was used to investigate the X-57 performance, without the distributed electric propulsion high-lift system operating. The unpowered X-57 performance is of interest to quantify if the X-57 can meet the cruise drag performance goal, and to document the lift performance of the very small wing at takeoff and landing conditions. The primary configurations investigated in this paper include the cruise configuration with no flap deflection, a takeoff configuration with a 10◦ flap deflection, and a landing configuration with a 30◦ flap deflection. The conditions for the cruise configuration were a flight unit Reynolds number of 1.32E+06 per foot, an altitude of 8000 feet, a Mach number of 0.233, and angles of attack from −2° to 24° . At the cruise lift coefficient of 0.7516, the computed drag coefficient is 0.05275. This computed drag is less than the drag coefficient of 0.05423 that is required to meet the X-57 airplane performance goal. However, the computational airplane is a completely smooth geometry and does not account for protuberance drag, nor the drag from steps and gaps in the actual X-57 airplane. Therefore, based upon the CFD drag calculation there is a 10-percent margin to account for some of the differences between the as-built metal fuselage and empennage construction, and the smooth computational geometry. The computed cruise drag also does not account for an induced drag reduction due to the wing-tip propellers and a drag reduction due to laminar flow achieved on the wing. The computed lift to drag ratio is 14.14 at the cruise lift coefficient of 0.7516, and the maximum computed lift to drag ratio is 15.8. The maximum lift coefficient for the cruise configuration was 2.13 at an angle of attack of 15°. The conditions for the takeoff configuration with a 10° flap deflection were a flight unit Reynolds number of 0.986E+06 per foot, an altitude of 2500 feet, a Mach number of 0.149, and angles of attack from −2° to 22°. The maximum lift coefficient for the takeoff configuration was 2.21 at an angle of attack of 16°. The conditions for the landing configuration with a 30° flap deflection were a flight unit Reynolds number of 0.922E+06 per foot, an altitude of 2500 feet, a Mach number of 0.139, and angles of attack from −2° to 24°. The maximum lift coefficient for the landing configuration was 2.58 and occurred at two angles of attack, 10° and 14°. Based on the unpowered maximum lift coefficient of 2.58 for the 30° flap deflection, along with computations of the distributed electric propulsion lift augmentation (not shown in this paper), the X-57 Maxwell is estimated to meet its powered landing goal of a maximum lift coefficient of 4.0.

X-57 Maxwell Airplane↗
NASA NTRS1947

Wind-Tunnel Investigation of the Effect of Tab Balance on Tab and Control-Surface Characteristics

An investigation was conducted to furnish data on the effect of tab balance on tab and control-surface characteristics. The airfoil tested had a modified NACA 65(1)-012 contour with a plain flap having a chord equal to 25 percent of the wing chord and with a tab having a chord equal to 25 percent of the flap chord and having several nose shapes and overhang lengths. The results of the investigation indicated that, in general, tab balance affected tab hinge-moment characteristics in much the same manner that flap balance affects flap hinge-moment characteristics. A moderate amount of tab balance did not seem to have any adverse effect on flap hinge-moment characteristics.

TABS, AILERON↗
DOE OSTI2023

Aero‐servo‐elastic co‐optimization of large wind turbine blades with distributed aerodynamic control devices

Abstract This work introduces automated wind turbine optimization techniques based on full aero‐servo‐elastic models and investigates the potential of trailing edge flaps to reduce the levelized cost of energy (LCOE) of wind turbines. The Wind Energy with Integrated Servo‐control (WEIS) framework is improved to conduct the presented research. Novel methods for the generic implementation and tuning of trailing edge flap devices and their controller are also introduced. Primary flap and controller parameters are optimized to demonstrate potential maximum blade tip deflection reductions of 21 % . Concurrent design optimization (i.e., co‐design) of a novel segmented wind turbine blade with trailing edge flaps and its controller is then conducted to demonstrate blade cost savings of 5 % . Additionally, rotor diameter co‐design optimization is demonstrated to reduce the LCOE by 1.3 % without significant load increases to the tower. These results demonstrate the efficacy of control co‐design optimization using trailing edge flaps, and the entirety of this work provides a foundation for numerous control co‐design‐oriented studies for distributed aerodynamic control devices.

17 WIND ENERGY↗