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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Effects of thrust reversing in ground proximity

The changes in stability and control characteristics encountered by a thrust reversing aircraft during its final approach, landing, and ground roll are described. These changes include a strong pitch-up accompanied by the loss of horizontal tail and aileron control effectiveness. The magnitude of reverser induced changes in ground effect are much larger than corresponding changes in free air. Some unexpected unsteady motions exhibited in wind tunnel by an aircraft model with reversers operating in ground proximity are also described. The cause of this oscillatory behavior was determined to be an unsteady interaction between the wall jets formed by impingement of reverser jets on the ground and the on-coming free stream. Time histories of rolling moments measured by the wind tunnel balance or support system were removed and frequencies were scaled by Strouhal number to full scale. Corrected time series were used to simulate the motion of a fighter aircraft with thrust reversers in ground effect. The simulation predicted large roll angles and nose down attitude at touchdown. Some phenomena of jet attachment to solid surfaces are discussed and areas for future research are recommended.

Joshi, P. B.↗
NASA NTRS1988

A review of technologies applicable to low-speed flight of high-performance aircraft investigated in the Langley 14- x 22-foot subsonic tunnel

An extensive research program has been underway at the NASA Langley Research Center to define and develop the technologies required for low-speed flight of high-performance aircraft. This 10-year program has placed emphasis on both short takeoff and landing (STOL) and short takeoff and vertical landing (STOVL) operations rather than on regular up and away flight. A series of NASA in-house as well as joint projects have studied various technologies including high lift, vectored thrust, thrust-induced lift, reversed thrust, an alternate method of providing trim and control, and ground effects. These technologies have been investigated on a number of configurations ranging from industry designs for advanced fighter aircraft to generic wing-canard research models. Test conditions have ranged from hover (or static) through transition to wing-borne flight at angles of attack from -5 to 40 deg at representative thrust coefficients.

Paulson, John W., Jr.↗
NASA NTRS1996

USAF Flight Test Investigation of Focused Sonic Booms: Project Have Bears

Supersonic operations from military aircraft generate sonic booms that can affect people, animals and structures. A substantial experimental data base exists on sonic booms for aircraft in steady flight and confidence in the predictive techniques has been established. All the focus sonic boom data that are in existence today were collected during the 60's and 70's as part of the information base to the US Supersonic Transport program and the French Jericho studies for the Concorde. These experiments formed the data base to develop sonic boom propagation and prediction theories for focusing. There is a renewed interest in high-speed transports for civilian application. Moreover, today's fighter aircraft have better performance capabilities, and supersonic flights ars more common during air combat maneuvers. Most of the existing data on focus booms are related to high-speed civil operations such as transitional linear accelerations and mild turns. However, military aircraft operating in training areas perform more drastic maneuvers such as dives and high-g turns. An update and confirmation of USAF prediction capabilities is required to demonstrate the ability to predict and control sonic boom impacts, especially those produced by air combat maneuvers.

Downing, Micah↗
NASA NTRS2003

Technical Evaluation Report, Part A - Vortex Flow and High Angle of Attack

A symposium entitled Vortex Flow and High Angle of Attack was held in Loen, Norway, from May 7 through May 11, 2001. The Applied Vehicle Technology (AVT) panel, under the auspices of the Research and Technology Organization (RTO), sponsored this symposium. Forty-eight papers, organized into nine sessions, addressed computational and experimental studies of vortex flows pertinent to both aircraft and maritime applications. The studies also ranged from fundamental fluids investigations to flight test results, and significant results were contributed from a broad range of countries. The principal emphasis of this symposium was on "the understanding and prediction of separation-induced vortex flows and their effects on military vehicle performance, stability, control, and structural design loads." It was further observed by the program committee that "separation- induced vortex flows are an important part of the design and off-design performance of conventional fighter aircraft and new conventional or unconventional manned or unmanned advanced vehicle designs (UAVs, manned aircraft, missiles, space planes, ground-based vehicles, and ships)." The nine sessions addressed the following topics: vortical flows on wings and bodies, experimental techniques for vortical flows, numerical simulations of vortical flows, vortex stability and breakdown, vortex flows in maritime applications, vortex interactions and control, vortex dynamics, flight testing, and vehicle design. The purpose of this paper is to provide brief reviews of these papers along with some synthesizing perspectives toward future vortex flow research opportunities. The paper includes the symposium program. (15 refs.)

VORTEX BREAKDOWN↗
NASA NTRS1991

IMPAC: An Integrated Methodology for Propulsion and Airframe Control

The National Aeronautics and Space Administration is actively involved in the development of enabling technologies that will lead towards aircraft with new/enhanced maneuver capabilities such as Short Take-Off Vertical Landing (STOVL) and high angle of attack performance. Because of the high degree of dynamic coupling between the airframe and propulsion systems of these types of aircraft, one key technology is the integration of the flight and propulsion control. The NASA Lewis Research Center approach to developing Integrated Flight Propulsion Control (IFPC) technologies is an in-house research program referred to as IMPAC (Integrated Methodology for Propulsion and Airframe Control). The goals of IMPAC are to develop a viable alternative to the existing integrated control design methodologies that will allow for improved system performance and simplicity of control law synthesis and implementation, and to demonstrate the applicability of the methodology to a supersonic STOVL fighter aircraft. Based on some preliminary control design studies that included evaluation of the existing methodologies, the IFPC design methodology that is emerging at the Lewis Research Center consists of considering the airframe and propulsion system as one integrated system for an initial centralized controller design and then partitioning the centralized controller into separate airframe and propulsion system subcontrollers to ease implementation and to set meaningful design requirements for detailed subsystem control design and evaluation. An overview of IMPAC is provided and detailed discussion of the various important design and evaluation steps in the methodology are included.

Garg, Sanjay↗
NASA NTRS1998

Lift Fan Nozzle for Joint Strike Fighter Tested in NASA Lewis' Powered Lift Rig

Under a nonreimbursable space act agreement between the NASA Lewis Research Center and the Allison Advanced Development Company, Allison tested a lift fan nozzle in Lewis' Powered Lift Rig. This test was in support of the Joint Strike Fighter program (formerly the Joint Advanced Strike Technology) sponsored by the Department of Defense, which will develop and field an affordable, multirole, next-generation, strike fighter aircraft for the Navy, Air Force, Marine Corps, and foreign allies. Allison, along with Pratt & Whitney Company, is part of the Lockheed Martin Corporation team that is scheduled to build a concept demonstrator aircraft by fiscal year 2001. The test was initiated in April and successfully completed in mid-July of 1997. Allison supplied a one-third-scale model of the lift fan nozzle, and Lewis provided the facility and the necessary support team. Various configurations, including pitching vectored angles ranging from 15deg forward to 60deg backward, were tested over a range of nozzle pressure ratios. Nozzle flow rates, thrust, and static pressures were measured for each of the configurations. Results from the test met the design requirements for the Joint Strike Fighter program and were in agreement with Allison's internal computational fluid dynamics (CFD) analyses. Data obtained from this test will also be used in the full-scale design of the lift fan system.

Lam, David W.↗
NASA NTRS1977

The optimal control frequency response problem in manual control

An optimal control frequency response problem is defined within the context of the optimal pilot model. The problem is designed to specify pilot model control frequencies reflective of important aircraft system properties, such as control feel system dynamics, airframe dynamics, and gust environment, as well as man machine properties, such as task and attention allocation. This is accomplished by determining a bounded set of control frequencies which minimize the total control cost. The bounds are given by zero and the neuromuscular control frequency response for each control actuator. This approach is fully adaptive, i.e., does not depend upon user entered estimates. An algorithm is developed to solve this optimal control frequency response problem. The algorithm is then applied to an attitude hold task for a bare airframe fighter aircraft case with interesting dynamic properties.

Harrington, W. W.↗
NASA NTRS1997

Robust, Decoupled, Flight Control Design with Rate Saturating Actuators

Techniques for the design of control systems for manually controlled, high-performance aircraft must provide the following: (1) multi-input, multi-output (MIMO) solutions, (2) acceptable handling qualities including no tendencies for pilot-induced oscillations, (3) a tractable approach for compensator design, (4) performance and stability robustness in the presence of significant plant uncertainty, and (5) performance and stability robustness in the presence actuator saturation (particularly rate saturation). A design technique built upon Quantitative Feedback Theory is offered as a candidate methodology which can provide flight control systems meeting these requirements, and do so over a considerable part of the flight envelope. An example utilizing a simplified model of a supermaneuverable fighter aircraft demonstrates the proposed design methodology.

Snell, S. A.↗
NASA NTRS1974

Digital adaptive flight control design using single stage model following indices

Simple mechanical linkages are often unable to cope with many control problems associated with high-performance aircraft. This has led to the development of digital fly-by-wire control systems and in particular digital adaptive controllers that can be efficiently adjusted during system operation. To this effect, a control law has been derived based upon the minimization of a single-stage weighted combination of control energy and the squared error between the states of a linear plant and model. This control logic is interfaced with an on-line weighted least-squares estimator and a Kalman state filter. The utility of the resultant control system is illustrated by its application to the linearized dynamics of a typical fighter aircraft.

Alag, G.↗
NASA NTRS1978

Aerodynamic characteristics of a jet sheet vortex generator

A configuration concept for augmenting the lift capability of low aspect ratio, thin wings, typically used on fighter aircraft was investigated. The fluid strake concept uses a jet sheet formed by blowing from a series of small orifices located in the side of the fuselage ahead of the wing to generate a stable vortex flow over the wing at high angle of attack. The effect of the location of the fluid strake relative to the wing was investigated for three different designs of the in-line orifices using a half-span model tested in a 7 by 10 foot low speed tunnel. Based on the results of the low speed test, a jet sheet producing module was incorporated into a NASA general research fighter model and tested in the Langley 7 by 10 foot high speed tunnel to determine the effectiveness of the fluid strake as a lift-enhancement device in the high-speed maneuver regime. Tests were conducted over a Mach number range from 0.3 to 0.8, with a jet momentum coefficient range from 0 to 0.24. Significant lift increments resulted at the higher angles of attack and drag polars were improved.

Ziegler, H.↗
NASA NTRS1986

Elliptic generation of composite three-dimensional grids about realistic aircraft

An elliptic method for generating composite grids about realistic aircraft is presented. A body-conforming grid is first generated about the entire aircraft by the solution of Poisson's differential equation. This grid has relatively coarse spacing, and it covers the entire physical domain. At boundary surfaces, cell size is controlled and cell skewness is nearly eliminated by inhomogeneous terms, which are found automatically by the program. Certain regions of the grid in which high gradients are expected, and which map into rectangular solids in the computational domain, are then designated for zonal refinement. Spacing in the zonal grids is reduced by adding points with a simple, algebraic scheme. Details of the grid-generation method are presented along with results of the present application, a wing/body configuration based on the F-16 fighter aircraft.

Sorenson, Reese L.↗
NASA NTRS1986

Elliptic generation of composite three-dimensional grids about realistic aircraft

An elliptic method for generating composite grids about realistic aircraft is presented. A body-conforming grid is first generated about the entire aircraft by the solution of Poisson's differential equation. This grid has relatively coarse spacing, and it covers the entire physical domain. At boundary surfaces, cell size is controlled and cell skewness is nearly eliminated by inhomogeneous terms, which are found automatically by the program. Certain regions of the grid in which high gradients are expected, and which map into rectangular solids in the computational domain, are then designated for zonal refinement. Spacing in the zonal grids is reduced by adding points with a simple, algebraic scheme. Details of the grid generation method are presented along with results of the present application, a wing-body configuration based on the F-16 fighter aircraft.

Sorenson, R. L.↗
NASA NTRS1992

A Regression Method for Comparing Launch Zone Ranges Derived From Different Sources

In modern fighter aircraft air-to-air missile launch acceptable ranges (LARs) are provided to the pilot for his launch decision. These LARs are calculated in the avionic software, using very simple algebraic equations and algorithms, that attempt to model complex missile behavior. Frequently these algorithms are revised to reflect inaccuracies discovered, when comparisons are made with missile six degree-of-freedom (6-DOF) simulations. The process of modifying this software has become expensive. In the past, these simple LAR generators were modified based only on subjective judgmental approach, with little regard to the missile range design limits. With reduced budgets a more rational methodology should be taken to assess whether or not simple fighter LAR generators need to be modified. Any LAR generator can produce large amounts of range (minimum and maximum) data over the board engagement limits of modem fighter and air-to-air missiles. This infinite population lends itself to some statistical methodology to access simple LAR accuracy and the need for software modification. The methodology should produce statistical parameters for use in the decision making process of modifying a simple fighter LAR generator. This paper presents a regression method, based on a missile range design limit philosophy. Acceptance criteria from the regression method are given that provide simple statistical parameters to aid in this decision making process.

Flight Testing↗
NASA NTRS1984

Approach and landing aerodynamic technologies for advanced STOL fighter configurations

Providing next-generation fighter aircraft with STOL capability will require effective high-lift systems, methods of providing adequate longitudinal trim and control, and thrust reversers compatible with advanced configurations. Recent wind-tunnel tests in the NASA Langley 4 x 7-m tunnel have provided the opportunity to investigate some promising technologies for providing this STOL capability. These tests include investigations of the high-lift contribution of spanwise blowing on a trailing-edge flap system, the trim capability of a blown high-lift canard, and the deceleration contribution of a thrust reverser for ground-roll reduction. This paper reviews the results of each investigation and examines the integration of these technologies on a common advanced-fighter configuration to determine their compatibility and the resulting overall performance of the configurations.

Banks, D. W.↗
NASA NTRS1994

Coupling equivalent plate and finite element formulations in multiple-method structural analyses

A coupled multiple-method analysis procedure for use late in conceptual design or early in preliminary design of aircraft structures is described. Using this method, aircraft wing structures are represented with equivalent plate models, and structural details such as engine/pylon structure, landing gear, or a 'stick' model of a fuselage are represented with beam finite element models. These two analysis methods are implemented in an integrated multiple-method formulation that involves the assembly and solution of a combined set of linear equations. The corresponding solution vector contains coefficients of the polynomials that describe the deflection of the wing and also the components of translations and rotations at the joints of the beam members. Two alternative approaches for coupling the methods are investigated; one using transition finite elements and the other using Lagrange multipliers. The coupled formulation is applied to the static analysis and vibration analysis of a conceptual design model of a fighter aircraft. The results from the coupled method are compared with corresponding results from an analysis in which the entire model is composed of finite elements.

VIBRATION ANALYSIS↗
NASA NTRS1999

Modeling of Longitudinal Unsteady Aerodynamics of a Wing-Tail Combination

Aerodynamic equations for the longitudinal motion of an aircraft with a horizontal tail were developed. In this development emphasis was given on obtaining model structure suitable for model identification from experimental data. The resulting aerodynamic models included unsteady effects in the form of linear indicial functions. These functions represented responses in the lift on the wing and tail alone, and interference between those two lifting surfaces. The effect of the wing on the tail was formulated for two different expressions concerning the downwash angle at the tail. The first expression used the Cowley-Glauert approximation known-as "lag-in-downwash," the second took into account growth of the wing circulation and delay in the development of the lift on the tail. Both approaches were demonstrated in two examples using the geometry of a fighter aircraft and a large transport. It was shown that the differences in the two downwash formulations would increase for an aircraft with long tail arm performing low-speed, rapid maneuvers.

Klein, Vladislav↗
NASA NTRS1995

Propulsion Controlled Aircraft design and development

This paper describes the design, development, and ground testing of the propulsion controlled aircraft (PCA) flight control system. A backup flight control system which uses only engine thrust, the PCA system utilizes collective and differential thrust changes to steer an aircraft that experiences partial or complete failure of the hydraulically actuated control surfaces. The objective of the program was to investigate, in flight, the throttles-only control capability of the F-15, using manual control, and also an augmented PCA mode in which computer-controlled thrust was used for flight control. The objective included PCA operation in up-and-away flight and, if performance was adequate, a secondary objective to make actual PCA landings. The PCA design began with a feasibility study which evaluated many control law designs. The study was done using off-line control analysis, simulation, and on-line manned flight simulator tests. Control laws, cockpit displays, and cockpit controls were evaluated by NASA test pilots. A flight test baseline configuration was selected based on projected flight performance, applicability to transport and fighter aircraft, and funding costs. During the PCA software and hardware development, the initial design was updated as data became available from throttle-only flight experiments conducted by NASA on the F-15. This information showed basic airframe characteristics that were not observed in the F-15 flight simulator and resulted in several design changes. After the primary objectives of the PCA flight testing were accomplished, additional PCA modes of operation were developed and implemented. The evolution of the PCA system from the initial feasibility study, control law design, simulation, hardware-in-the-loop tests, pilot-in-the-loop tests, and ground tests is presented.

Wells, Edward A.↗
NASA NTRS1988

Effects of symmetric and asymmetric modes on transonic aeroelastic characteristics of full-span wing-body configurations

To accurately study the transonic aeroelastic characteristics, it is important to model the full aircraft configuration, including asymmetry. Recently, an accurate method of computing unsteady transonic flows on full-span wing-body configurations was developed using the transonic potential flow theory. In this work, the method is further developed to account for the aeroelasticity of full-span wing-body configurations. This is accomplished by simultaneously integrating the unsteady aerodynamic forces and modal structural equations of the wing-body configurations. To validate the method, aeroelastic computations are made for a wing-body configuration with a rectangular wing. The aeroelastic responses of this configuration are correlated with the responses of a similar isolated wing. The comparisons are favorable. Aeroelastic computations associated with symmetric and asymmetric modes are also made to study the influence of modal asymmetry on responses. This new development is further illustrated by computing aeroelastic characteristics of a typical fighter aircraft. The results from this study will be useful in accurately computing the transonic flutter boundaries of aircraft, including those associated with asymmetric modes.

Guruswamy, Guru P.↗