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At least 145 records · Page 8

Fusion interfaces for tactical environments: An application of virtual reality technology

The term Fusion Interface is defined as a class of interface which integrally incorporates both virtual and nonvirtual concepts and devices across the visual, auditory, and haptic sensory modalities. A fusion interface is a multisensory virtually-augmented synthetic environment. A new facility has been developed within the Human Engineering Division of the Armstrong Laboratory dedicated to exploratory development of fusion interface concepts. This new facility, the Fusion Interfaces for Tactical Environments (FITE) Facility is a specialized flight simulator enabling efficient concept development through rapid prototyping and direct experience of new fusion concepts. The FITE Facility also supports evaluation of fusion concepts by operation fighter pilots in an air combat environment. The facility is utilized by a multidisciplinary design team composed of human factors engineers, electronics engineers, computer scientists, experimental psychologists, and oeprational pilots. The FITE computational architecture is composed of twenty-five 80486-based microcomputers operating in real-time. The microcomputers generate out-the-window visuals, in-cockpit and head-mounted visuals, localized auditory presentations, haptic displays on the stick and rudder pedals, as well as executing weapons models, aerodynamic models, and threat models.

Haas, Michael W.↗

A comprehensive analytical model of rotorcraft aerodynamics and dynamics. Part 1: Analysis development

Structural, inertia, and aerodynamic models were combined to form a comprehensive model of rotor aerodynamics and dynamics that is applicable to a wide range of problems and a wide class of vehicles. A digital computer program is used to calculate rotor performance, loads, and noise; helicopter vibration and gust response; flight dynamics and handling qualities; and system aeroelastic stability. The analysis is intended for use in the design, testing, and evaluation of rotors and rotorcraft, and to be a basis for further development of rotary wing theories.

Johnson, W.↗

Aeroelastic forced response analysis of turbomachinery

An introduction is given to the research activity that is underway to enable the prediction of turbomachinery aeroelastic forced response. An effort is being made to assemble a computer program (FREPS) which incorporates the aeroelastic structural models, unsteady aerodynamic models, and forcing function models. The structural and aerodynamic models are currently well developed. The forcing function models are at a primitive level. A significant activity has begun to identify the forcing functions due to stator-rotor aerodynamic interaction.

Smith, Todd E.↗

Aerodynamic Characteristics of a 45 Deg Swept Wing Fighter Airplane Model and Aerodynamic Loads on Adjacent Stores and Missiles at Mach Numbers of 1.57, 1.87, 2.16, and 2.53

An investigation was performed in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a model of a 45 deg swept-wing fighter airplane, and to determine the loads on attached stores and detached missiles in the presence of the model. Also included was a determination of aileron-spoiler effectiveness, aileron hinge moments, and the effects of wing modifications on model aerodynamic characteristics. Tests were performed at Mach numbers of 1.57, 1.87, 2.16, and 2.53. The Reynolds numbers for the tests, based on the mean aerodynamic chord of the wing, varied from about 0.9 x 10(exp 6) to 5 x 10(exp 6). The results are presented with minimum analysis.

Oehman, W. I.↗

Aerodynamic Characteristics of a 45 Degree Swept-wing Fighter-Airplane Model and Aerodynamic Loads on Adjacent Stores and Missiles at Mach Numbers of 1.57, 1.87, 2.16, and 2.53

An investigation was performed in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a model of a 450 swept-wing fighter airplane, and to determine the loads on attached stores and detached missiles in the presence of the model. Also included was a determination of aileron-spoiler effectiveness, aileron hinge moments, and the effects of wing modifications on model aerodynamic characteristics. Tests were performed at Mach numbers of 1.57, 1.87, 2.16, and 2.53. The Reynolds numbers for the tests, based on the mean aerodynamic chord of the wing, varied from about 0.9 x 10(exp 6) to 5 x 10(exp 6). The results are presented with minimum analysis.

Oehman, Waldo I↗

Rarefied-flow shuttle aerodynamics flight model

A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low frequency, low level acceleration measurements are made during atmospheric reentry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP reentry mission data sets spanning a period of about 10 years have been processed. The HiRAP derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as functions of angle-of-attack, body flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle-of-attack are also presented, along with flight derived rarefied-flow transition bridging formulas. Comparisons are made between the aerodynamics model and data from the Orbiter Operational Aerodynamic Design Data Book, and applicable computer simulations and wind-tunnel data.

Blanchard, Robert C.↗

Rarefield-Flow Shuttle Aerodynamics Flight Model

A model of the Shuttle Orbiter rarefied-flow aerodynamic force coefficients has been derived from the ratio of flight acceleration measurements. The in-situ, low-frequency (less than 1Hz), low-level (approximately 1 x 10(exp -6) g) acceleration measurements are made during atmospheric re-entry. The experiment equipment designed and used for this task is the High Resolution Accelerometer Package (HiRAP), one of the sensor packages in the Orbiter Experiments Program. To date, 12 HiRAP re-entry mission data sets spanning a period of about 10 years have been processed. The HiRAP-derived aerodynamics model is described in detail. The model includes normal and axial hypersonic continuum coefficient equations as function of angle of attack, body-flap deflection, and elevon deflection. Normal and axial free molecule flow coefficient equations as a function of angle of attack are also presented, along with flight-derived rarefied-flow transition bridging formulae. Comparisons are made between the aerodynamics model, data from the latest Orbiter Operational Aerodynamic Design Data Book, applicable computer simulations, and wind-tunnel data.

Blanchard, Robert C.↗

Progress and challenges in modeling turbulent aerodynamic flows

Progress in modeling external aerodynamic flows achieved by using computations and experiments designed to guide turbulence modeling is presented. The computational procedures emphasize utilization of the Reynolds-averaged Navier-Stokes equations and various statistical modeling approaches. Developments for including the influence of compressibility are provided; they point up some of the complexities involved in modeling high-speed flows. Examples of complementary studies that provide the status, limitations, and future challenges of modeling for transonic, supersonic, and hypersonic flows are given.

Marvin, Joseph G.↗

New Flutter Analysis Technique for CFD-based Unsteady Aeroelasticity

This paper presents a flutter analysis technique for the transonic flight regime. The technique uses an iterative approach to determine the critical dynamic pressure for a given mach number. Unlike other CFD-based flutter analysis methods, each iteration solves for the critical dynamic pressure and uses this value in subsequent iterations until the value converges. This process reduces the iterations required to determine the critical dynamic pressure. To improve the accuracy of the analysis, the technique employs a known structural model, leaving only the aerodynamic model as the unknown. The aerodynamic model is estimated using unsteady aeroelastic CFD analysis combined with a parameter estimation routine. The technique executes as follows. The known structural model is represented as a finite element model. Modal analysis determines the frequencies and mode shapes for the structural model. At a given mach number and dynamic pressure, the unsteady CFD analysis is performed. The output time history of the surface pressure is converted to a nodal aerodynamic force vector. The forces are then normalized by the given dynamic pressure. A multi-input multi-output parameter estimation software, ERA, estimates the aerodynamic model through the use of time histories of nodal aerodynamic forces and structural deformations. The critical dynamic pressure is then calculated using the known structural model and the estimated aerodynamic model. This output is used as the dynamic pressure in subsequent iterations until the critical dynamic pressure is determined. This technique is demonstrated on the Aerostructures Test Wing-2 model at NASA's Dryden Flight Research Center.

Pak, Chan-gi↗

Roll-yaw control at high angle of attack by forebody tangential blowing

The feasibility of using forebody tangential blowing to control the roll-yaw motion of a wind tunnel model is experimentally demonstrated. An unsteady model of the aerodynamics is developed based on the fundamental physics of the flow. Data from dynamic experiments is used to validate the aerodynamic model. A unique apparatus is designed and built that allows the wind tunnel model two degrees of freedom, roll and yaw. Dynamic experiments conducted at 45 degrees angle of attack reveal the system to be unstable. The natural motion is divergent. The aerodynamic model is incorporated into the equations of motion of the system and used for the design of closed loop control laws that make the system stable. These laws are proven through dynamic experiments in the wind tunnel using blowing as the only actuator. It is shown that asymmetric blowing is a highly non-linear effector that can be linearized by superimposing symmetric blowing. The effects of forebody tangential blowing and roll and yaw angles on the flow structure are determined through flow visualization experiments. The transient response of roll and yaw moments to a step input blowing are determined. Differences on the roll and yaw moment dependence on blowing are explained based on the physics of the phenomena.

Pedreiro, N.↗

Roll-Yaw control at high angle of attack by forebody tangential blowing

The feasibility of using forebody tangential blowing to control the roll-yaw motion of a wind tunnel model is experimentally demonstrated. An unsteady model of the aerodynamics is developed based on the fundamental physics of the flow. Data from dynamic experiments is used to validate the aerodynamic model. A unique apparatus is designed and built that allows the wind tunnel model two degrees of freedom, roll and yaw. Dynamic experiments conducted at 45 degrees angle of attack reveal the system to be unstable. The natural motion is divergent. The aerodynamic model is incorporated into the equations of motion of the system and used for the design of closed loop control laws that make the system stable. These laws are proven through dynamic experiments in the wind tunnel using blowing as the only actuator. It is shown that asymmetric blowing is a highly non-linear effector that can be linearized by superimposing symmetric blowing. The effects of forebody tangential blowing and roll and yaw angles on the flow structure are determined through flow visualization experiments. The transient response of roll and yaw moments to a step input blowing are determined. Differences on the roll and yaw moment dependence on blowing are explained based on the physics of the phenomena.

Pedreiro, N.↗

Aerodynamic Loading Characteristics at Mach Numbers from 0.80 to 1.20 of a 1/10-Scale Three-Stage Scout Model

Aerodynamic loads results have been obtained in the Langley 8-foot transonic pressure tunnel at Mach numbers from 0.80 to 1.20 for a 1/10-scale model of the upper three stages of the Scout vehicle. Tests were conducted through an angle-of-attack range from -8 deg to 8 deg at an average test Reynolds number per foot of about 4.0 x 10(exp 6). Results indicated that the peak negative pressures associated with expansion corners at the nose and transition flare exhibit sizeable variations which occur over a relatively small Mach number range. The magnitude of the variations may cause the critical local loading condition for the full-scale vehicle to occur at a Mach number considerably lower than that at which the maximum dynamic pressure occurs in flight. The addition of protuberances simulating antennas and wiring conduits had slight, localized effects. The lift carryover from the nose and transition flare on the cylindrical portions of the model generally increased with an increase in Mach number.

Kelly, Thomas C.↗

Aerodynamic mathematical modeling - basic concepts

This lecture is a review of the basic concepts involved In the mathematical modeling of the aerodynamic response of an aircraft to arbitrary maneuvers. The review leads from Bryan's original formulation through the concepts of linear aerodynamic indicial functions and superposition to the extension of these concepts into the nonlinear regime. The nonlinear generalization yields a form for the aerodynamic response that can be built up from the responses to a limited number of well-defined characteristic motions, reproducible in principle either in wind-tunnel experiments or flow-field computations. A further generalization leads to a torn, accommodating the discontinuous and double-valued behavior characteristic of hysteresis in the steady-state aerodynamic response

Murray Tobak↗

Turbulence modeling for aerodynamic flows

The state of statistical turbulence modeling is reviewed with emphasis on compressible aerodynamic flows. A three-stage modeling process is suggested to act as a frame-work for improving turbulence models for complex flows. This process is then used to guide assessments of the comparative performance of various levels of turbulence models for several examples of flow fields of aerodynamic interest. The examples cover flows in two and three dimensions and with a broad range of Mach numbers. In addition to considerations of individual flow fields, some generalizations are drawn from the aggregate behavior of the comparisons.

Rubesin, Morris W.↗

User Selection Criteria of Airspace Designs in Flexible Airspace Management

A method for identifying global aerodynamic models from flight data in an efficient manner is explained and demonstrated. A novel experiment design technique was used to obtain dynamic flight data over a range of flight conditions with a single flight maneuver. Multivariate polynomials and polynomial splines were used with orthogonalization techniques and statistical modeling metrics to synthesize global nonlinear aerodynamic models directly and completely from flight data alone. Simulation data and flight data from a subscale twin-engine jet transport aircraft were used to demonstrate the techniques. Results showed that global multivariate nonlinear aerodynamic dependencies could be accurately identified using flight data from a single maneuver. Flight-derived global aerodynamic model structures, model parameter estimates, and associated uncertainties were provided for all six nondimensional force and moment coefficients for the test aircraft. These models were combined with a propulsion model identified from engine ground test data to produce a high-fidelity nonlinear flight simulation very efficiently. Prediction testing using a multi-axis maneuver showed that the identified global model accurately predicted aircraft responses.

Lee, Hwasoo E.↗

An Investigation of the Effects of a Vortex-Generator Configuration on the Aerodynamic Characteristics of a 1/4-Scale Model of the X-1 Airplane (10-Percent-Thick Wing)

An investigation of a vortex-generator configuration on the wings of a l/4-scale model of the X-1 airplane having a 10-percent-thick wing was conducted in the Langley 16-foot transonic tunnel. The effect of the vortex generators was determined by comparing the model aerodynamic characteristics, wing-pressure distributions, and wing-wake patterns for model configurations with and without vortex generators on the wings. Results are presented from tests at 0.1 increments in Mach number from about 0.69 to 0.99, at Reynolds numbers of about 4.1 x 10(exp 6) to 4.7 x 10(exp 6), and through an angle-of-attack range up to 1.5 deg at lower speeds and up to 5 deg at the highest speed. In general, little difference in the aerodynamic characteristics was observed, except at a Mach number of 0.90 where a rearward movement of the shock on the upper surface of the wing with the vortex generators installed resulted in less separation.

Runckel, Jack F.↗