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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 433 records · Page 24

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NASA NTRS2005

Active Tailoring of Lift Distribution to Enhance Cruise Performance

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.

Flamm, Jeffrey D.↗

NASA NTRS1977

Forward flight effects on EBF noise

Forward flight effects on local mean velocity and turbulence velocity profiles, surface pressure spectra, and far field acoustic pressure spectra were measured for a simple externally blown flap (EBF). Both upper-surface-blowing and under-the-wing configurations were tested. Ratio of acoustic wind tunnel velocity to nozzle exhaust velocity was varied from 0 to 3/8 in steps of 1/8. A method was determined for predicting forward flight effects on surface-radiated noise. This noise is decreased in amplitude and shifted to higher frequency relative to data obtained at zero flight speed. Predictions are validated by comparisons with published NASA, Boeing, and Lockheed data.

Fink, M. R.↗

NASA NTRS1996

A Passive Cavity Concept for Improving the Off-Design Performance of Fixed-Geometry Exhaust Nozzles

An investigation was conducted in the model preparation area of the Langley 16-Foot Transonic Tunnel to study a passive cavity concept for improving the off-design performance of fixed-geometry exhaust nozzles. Passive cavity ventilation (through a porous surface) was applied to divergent flap surfaces and tested at static conditions in a sub-scale, nonaxisymmetric, convergent-divergent nozzle. As part of a comprehensive investigation, force, moment and pressure measurements were taken and focusing schlieren flow visualization was obtained for a baseline configuration and D passive cavity configurations. All tests were conducted with no external flow and high-pressure air was used to simulate jet-exhaust flow at nozzle pressure ratios from 1.25 to approximately 9.50. Results indicate that baseline nozzle performance was dominated by unstable shock-induced boundary-layer separation at off-design conditions, which came about through the natural tendency of overexpanded exhaust flow to satisfy conservation requirements by detaching from the nozzle divergent flaps. Passive cavity ventilation added the ability to control off-design separation in the nozzle by either alleviating separation or encouraging stable separation of the exhaust flow. Separation alleviation offers potential for installed nozzle performance benefits by reducing drag at forward flight speeds, even though it may reduce off-design static thrust efficiency as much as 3.2 percent. Encouraging stable separation of the exhaust flow offers significant performance improvements at static, low NPR and low Mach number flight conditions by improving off-design static thrust efficiency as much as 2.8 percent. By designing a fixed-geometry nozzle with fully porous divergent flaps, where both cavity location and percent open porosity of the flaps could be varied, passive flow control would make it possible to improve off-design nozzle performance across a wide operating range. In addition, the ability to encourage separation on one flap while alleviating it on the other makes it possible to generate thrust vectoring in the nozzle through passive flow control.

Asbury, Scott C.↗

NASA NTRS1976

Study of an aircraft decoupled longitudinal control system for approach and landing

A series of ground-based and in-flight simulation studies of the application of steady-state decoupled longitudinal controls to a short take-off and landing (STOL) transport have been made. The externally blown flap STOL was selected for study because it was considered to be a worst-case situation from the control viewpoint. The decoupled longitudinal control system used constant prefilter and feedback gains to provide independent control of flight-path angle, pitch angle, and forward velocity during landing approach. The decoupled controls were compared to a more conventional stability augmentation system. The pilots were enthusiastic about the decoupled controls; the pilot workload was reduced and the landing performance significantly improved. The benefit of the decoupled controls was more dramatic during in-flight simulation using a variable stability airplane than was the case with either fixed- or moving-base, ground-based simulators.

Miller, G. K., Jr.↗

NASA NTRS1948

An Investigation of Aileron Oscillations at Transonic Speeds on NACA 23012 and NACA 65-212 Airfoils by the Wing-flow Method

An investigation is being conducted to determine the feasibility of studying aileron buzz by means of the wing-flow method. Two semispan models which had an aspect ratio of 6 and a taper ratio of 2 with quarter-chord half-span mass-balanced ailerons have been used. One had an NACA 23012 airfoil section and the second, an NACA 65-212 airfoil section. The ailerons on both models were subject to buzz over a small range of Mach number near 0.9. Data obtained by wing-flow testing agreed reasonably well with full-scale flight results.

MACH NUMBER EFFECTS - COMPLETE WINGS↗

NASA NTRS1946

Flight Investigation at High Speeds of Profile Drag of Wing of a P-47D Airplane Having Production Surfaces Covered with Camouflage Paint

Wing section outboard of flap was tested by wake surveys in Mach range of 0.25 - 0.78 and lift coefficient range 0.06 - 0.69. Results indicated that minimum profile-drag coefficient of 0.0097 was attained for lift coefficients from 0.16 to 0.25 at Mach less than 0.67. Below Mach number at which compressibility shock occurred, variations in Mach of 0.2 had negligible effect on profile drag coefficient. Shock was not evident until critical Mach was exceeded by 0.025.

DRAG - WINGS EFFECT OF PAINTS↗

NASA NTRS1987

Effect of helicopter blade dynamics on blade aerodynamic and structural loads

The effect of rotor blade dynamics on aerodynamic and structural loads is examined for a conventional, main-rotor helicopter using a comprehensive rotorcraft analysis (CAMRAD) and flight-test data. The impact of blade dynamics on blade section lift-coefficient time histories is studied by comparing predictions from a rigid-blade analysis and an elastic-blade analysis with helicopter flight test data. The elastic blade analysis better predicts high-frequency behavior of section lift. In addition, components of the blade angle of attack such as elastic blade twist, blade flap rate, blade slope velocity, and inflow are examined as a function of blade mode. Elastic blade motion changed blade angle of attack by a few tenths of a degree, and up to the sixth rotor harmonic. A similar study of the influence of blade dynamics on bending and torsion moments was also conducted. A correlation study comparing predictions from several elastic-blade analyses with flight-test data revealed that an elastic-blade model consisting of only three elastic bending modes (first and second flap and first lag), and two elastic torsion modes was sufficient for good correlation.

Heffernan, Ruth M.↗

NASA NTRS1993

In-flight pressure distributions and skin-friction measurements on a subsonic transport high-lift wing section

Flight experiments are being conducted as part of a multiphased subsonic transport high-lift research program for correlation with wind-tunnel and computational results. The NASA Langley Transport Systems Research Vehicle (B737-100 aircraft) is used to obtain in-flight flow characteristics at full-scale Reynolds numbers to contribute to the understanding of 3-D high-lift, multi-element flows including attachment-line transition and relaminarization, confluent boundary-layer development, and flow separation characteristics. Flight test results of pressure distributions and skin friction measurements were obtained for a full-chord wing section including the slat, main-wing, and triple-slotted, Fowler flap elements. Test conditions included a range of flap deflections, chord Reynolds numbers (10 to 21 million), and Mach numbers (0.16 to 0.40). Pressure distributions were obtained at 144 chordwise locations of a wing section (53-percent wing span) using thin pressure belts over the slat, main-wing, and flap elements. Flow characteristics observed in the chordwise pressure distributions included leading-edge regions of high subsonic flows, leading-edge attachment-line locations, slat and main-wing cove-flow separation and reattachment, and trailing-edge flap separation. In addition to the pressure distributions, limited skin-friction measurements were made using Preston-tube probes. Preston-tube measurements on the slat upper surface suggested relaminarization of the turbulent flow introduced by the pressure belt on the slat leading-edge surface when the slat attachment line was laminar. Computational analysis of the in-flight pressure measurements using two-dimensional, viscous multielement methods modified with simple-sweep theory showed reasonable agreement. However, overprediction of the pressures on the flap elements suggests a need for better detailed measurements and improved modeling of confluent boundary layers as well as inclusion of three-dimensional viscous effects in the analysis.

Yip, Long P.↗

NASA NTRS1998

Simulation-Based Analysis of Reentry Dynamics for the Sharp Atmospheric Entry Vehicle

This thesis describes the analysis of the reentry dynamics of a high-performance lifting atmospheric entry vehicle through numerical simulation tools. The vehicle, named SHARP, is currently being developed by the Thermal Protection Materials and Systems branch of NASA Ames Research Center, Moffett Field, California. The goal of this project is to provide insight into trajectory tradeoffs and vehicle dynamics using simulation tools that are powerful, flexible, user-friendly and inexpensive. Implemented Using MATLAB and SIMULINK, these tools are developed with an eye towards further use in the conceptual design of the SHARP vehicle's trajectory and flight control systems. A trajectory simulator is used to quantify the entry capabilities of the vehicle subject to various operational constraints. Using an aerodynamic database computed by NASA and a model of the earth, the simulator generates the vehicle trajectory in three-dimensional space based on aerodynamic angle inputs. Requirements for entry along the SHARP aerothermal performance constraint are evaluated for different control strategies. Effect of vehicle mass on entry parameters is investigated, and the cross range capability of the vehicle is evaluated. Trajectory results are presented and interpreted. A six degree of freedom simulator builds on the trajectory simulator and provides attitude simulation for future entry controls development. A Newtonian aerodynamic model including control surfaces and a mass model are developed. A visualization tool for interpreting simulation results is described. Control surfaces are roughly sized. A simple controller is developed to fly the vehicle along its aerothermal performance constraint using aerodynamic flaps for control. This end-to-end demonstration proves the suitability of the 6-DOF simulator for future flight control system development. Finally, issues surrounding real-time simulation with hardware in the loop are discussed.

Tillier, Clemens Emmanuel↗

NASA NTRS1973

Maneuver and buffet characteristics of fighter aircraft

Recent research efforts in the improvement of the maneuverability of fighter aircraft in the high-subsonic and transonic speed range are reviewed with emphasis on the factors affecting aerodynamic boundaries, such as maximum obtainable lift, buffet onset, pitchup, wing rock, and nose slice. The investigations were made using a general research configuration which encompassed a systematic matrix of wing-design parameters. These results illustrated the sensitivity of section and planform geometry to a selected design point. The incorporation of variable-geometry wing devices in the form of flaps or leading-edge slats was shown to provide controlled flow over a wide range of flight conditions and substantial improvements in maneuver capabilities. Additional studies indicated that the blending of a highly swept maneuver strake with an efficient, moderately swept wing offers a promising approach for improving maneuver characteristics at high angles of attack without excessive penalties in structural weight.

Ray, E. J.↗

NASA NTRS1974

On the use of first order rotor dynamics in multiblade coordinates

This paper is directed to the question of how to represent most efficiently rotor/body coupling in a linear flight dynamics analysis. Rigid body pitch, roll and vertical motions are considered for the rotor/body coupling studies. Flapping stability limits, eigenvalues, transient responses to control step inputs, to step gusts and to random gusts are determined for a hypothetical hingeless compound helicopter operating up to .8 advance ratio. Data are obtained for the basic helicopter and for the craft with two simple control feedback systems. While complete periodic system modeling is necessary for determining flapping stability limits and vibrations, constant system modeling using first order dynamics in each of the multiblade rotor coordinates was found to be adequate for rotor-craft stability and response computations.

Hohenemser, K. H.↗

NASA NTRS1976

On the status of V/STOL flight

Basic principles relating to the accomplishment of V/STOL flight are reviewed as they pertain to current prototype developments. Particular consideration is given to the jet flap, flow augmentation and circulation control separately and in combination. Configurations as the augmentor wing, upper-surface blown flaps, externally blown flaps and the circulation-controlled rotor are also examined.

Mccormick, B. W.↗

NASA NTRS1979

Manufacturing processes for fabricating graphite/PMR 15 polyimide structural elements

Investigations were conducted to obtain commercially available graphite/PMR-15 polyimide prepreg, develop an autoclave manufacturing process, and demonstrate the process by manufacturing structural elements. Controls were established on polymer, prepreg, composite fabrication, and quality assurance, Successful material quality control and processes were demonstrated by fabricating major structural elements including flat laminates, hat sections, I beam sections, honeycomb sandwich structures, and molded graphite reinforced fittings. Successful fabrication of structural elements and simulated section of the space shuttle aft body flap shows that the graphite/PMR-15 polyimide system and the developed processes are ready for further evaluation in flight test hardware.

Sheppard, C. H.↗

NASA NTRS1986

Symbolic generation of elastic rotor blade equations using a FORTRAN processor and numerical study on dynamic inflow effects on the stability of helicopter rotors

The process of performing an automated stability analysis for an elastic-bladed helicopter rotor is discussed. A symbolic manipulation program, written in FORTRAN, is used to aid in the derivation of the governing equations of motion for the rotor. The blades undergo coupled bending and torsional deformations. Two-dimensional quasi-steady aerodynamics below stall are used. Although reversed flow effects are neglected, unsteady effects, modeled as dynamic inflow are included. Using a Lagrangian approach, the governing equations are derived in generalized coordinates using the symbolic program. The program generates the steady and perturbed equations and writes into subroutines to be called by numerical routines. The symbolic program can operate on both expressions and matrices. For the case of hovering flight, the blade and dynamic inflow equations are converted to equations in a multiblade coordinate system by rearranging the coefficients of the equations. For the case of forward flight, the multiblade equations are obtained through the symbolic program. The final multiblade equations are capable of accommodating any number of elastic blade modes. The computer implementation of this procedure consists of three stages: (1) the symbolic derivation of equations; (2) the coding of the equations into subroutines; and (3) the numerical study after identifying mass, damping, and stiffness coefficients. Damping results are presented in hover and in forward flight with and without dynamic inflow effects for various rotor blade models, including rigid blade lag-flap, elastic flap-lag, flap-lag-torsion, and quasi-static torsion. Results from dynamic inflow effects which are obtained from a lift deficiency function for a quasi-static inflow model in hover are also presented.

Reddy, T. S. R.↗

NASA NTRS1982

Correlation of flight effects on centerline velocity decay for cold-flow acoustically excited jets

Acoustic excitation can influence the large scale shear layer structure of jet flow, thereby causing the jet centerline velocity to decay more rapidly. this phenomena has numerous practical applications, such as the reduction of jet/flap impingement velocity in order to reduce flap structural loads for under-the-wing STOL aircraft concepts. Cold flow centerline velocity decay data obtained with acoustic excitation of the jet exhaust flow are correlated with and without flight effects. Initially, static data are correlated in terms of an acoustic parameter involving the threshold sound level needed to initiate acoustic excitation of the jet plume. It is shown that for a given flight speed, the same acoustic excitation parameter as that developed for the static condition is valid. Finally, an aerodynamic flight effects parameter is included to correlate the static and flight centerline velocity decay data with and without acoustic excitation.

Vonglahn, U. H.↗

NASA NTRS1947

Development of Inboard Nacelle for the XB-36 Airplane

A series of investigations of several 1/14-scale models of an inboard nacelle for the XB-36 airplane was made in the Langley two-dimensional low-turbulence tunnels. The purpose of these investigations was to develop a low-drag wing-nacelle pusher combination which incorporated an internal air-flow system. As a result of these investigations, a nacelle was developed which had external drag coefficients considerably lower than the original basic form with the external nacelle drag approximately one-half to two-thirds of those of conventional tractor designs. The largest reductions in drag resulted from sealing the gaps between the wing flaps and nacelle, reducing the thickness of the nacelle training-edge lip, and bringing the under-wing air inlet to the wing leading edge. It was found that without the engine cooling fan adequate cooling air would be available for all conditions of flight except for cruise and climb at 40,000 feet. Sufficient oil cooling at an altitude of 40,000 feet may be obtained by the use of flap-type exit doors.

Nuber, Robert J.↗

NASA NTRS1981

Recent Progress in V/STOL Aircraft Technology

Results from wind tunnel and flight tests investigations for V/STOL aircraft are reviewed. Primary emphasis is given to technical results relating to three types of subsonic aircraft: a quiet STOL aircraft; a tilt rotor aircraft; and a turbofan V/STOL aircraft. Comparison and correlation between theoretical and experimental results and between wind tunnel and flight test results, is made. The quiet STOL aircraft technology results are primarily those derived from the NASA/Boeing Quiet Short Haul Technology (QSRA) program. The QSRA aircraft uses an upper surface blown flap and develops a usable engine-out landing approach lift coefficient of 5.5 and landing distances less than 1,000 ft. The tilt rotor aircraft technology results are those obtained from the NASA/Army/Navy/Bell (XV-15-TRRA) aircraft flight investigations. The TRRA is a twin rotor research aircraft capable of vertical takeoff and landing and cruise speeds of 300 knots. The turbofan V/STOL aircraft technology results are from static ground facility and wind tunnel investigations of a NASA/NAVY/Grumman full scale lift/cruise fan aircraft model, which features two tilting nacelles with TF-34 engines.

Roberts, L.↗

NASA NTRS2013

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

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.

flight optimization↗