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NASA NTRSDate not supplied

FUN3D Analyses in Support of the 1st AIAA Stability and Control Prediction Workshop

The 1st AIAA Stability and Control Prediction Workshop was created to establish best practices for the prediction of stability & control derivatives using computational fluid dynamics and assess the limitations of these computational methods when those best practices are applied. The inaugural workshop considers the ONERA version of the NASA/Boeing Common Research Model(CRM), which includes the wing, body, horizontal tail, and a vertical tail designed by ONERA. Wind tunnel tests have been conducted for this configuration with longitudinal tests having been previously published, in addition to unpublished data at small sideslip angles that will serve as ‘blind’ data for workshop data comparisons. Participants were provided a ‘family’ of unstructured grids for the full-span ONERA CRM model with the wind-tunnel sting included. This family of mixed-element grids consists of 5 levels of refinement (tiny, coarse, medium, fine, and extra fine) with surface and volume mesh scaling, resulting in a size range of 14.6 to 53.4 million nodes. In addition, a medium refinement mesh has been provided for the ONERA CRM configuration without a sting to evaluate the sting’s impact on static longitudinal stability characteristics. In addition to these workshop-provided grids, the present work also considers an equivalent ‘family’ of computational grids generated using Heldenmesh™, a rapid grid generation software by Helden Aerospace Corporation for creating high-quality, three-dimensional, mixed-element unstructured meshes. Because of the authors’ familiarity with this software, these additional grids were generated as a comparison to the workshop-provided grids and to better understand the implications of using volume-mirrored grids for stability and control predictions. The present work will contribute to the workshop with test case data generated using the NASA FUN3D code, a parallelized, unstructured, node-based, finite-volume discretization, Reynolds-averaged Navier-Stokes flow solver. Numerical simulations will be conducted using the Quadratic Constitutive Relationship (QCR) version of the Spalart-Allmaras (SA) turbulence model with negative turbulence variable provisions. Both steady and2nd-order, time-accurate simulation results are to be generated and compared for select test cases, as time permits, to investigate their impact on FUN3D predictions. The present work will consider the three primary workshop test cases: (1) grid convergence study, (2) Mach number effect on static stability, and (3) wind tunnel sting increments. Additionally, data will be provided for the two optional test cases, which include:(1) static stability derivative calculations and (2)sideslip angle sweeps. In each of the test cases, the vehicle is stationary, and the body is assumed to be rigid, where vehicle deformation has been accounted for in the model configuration geometry. For all test cases, longitudinal and lateral force and moment aerodynamic coefficients will be provided for the total configuration, in addition to a component-level breakdown that includes the port wing, starboard wing, fuselage, and tail.

CFD↗
NASA NTRS1999

Power Effects on High Lift, Stability and Control Characteristics of the TCA Model Tested in the LaRC 14 x 22 Ft Wind Tunnel

The TCA-2 wind-tunnel test was the second in a series of planned tests utilizing the 5% Technology Concept Airplane (TCA) model. Each of the tests was planned to utilize the unique capabilities of the NASA Langley 14'x22' and the NASA Ames 12' test facilities, in order to assess specific aspects of the high lift and stability and control characteristics of the TCA configuration. However, shortly after the completion of the TCA-1 test, an early projection of the Technology Configuration (TC) identified the need for several significant changes to the baseline TCA configuration. These changes were necessary in order to meet more stringent noise certification levels, as well as, to provide a means to control dynamic structural modes. The projected changes included a change to the outboard wing (increased aspect ratio and lower sweep) and a reconfiguration of the longitudinal control surfaces to include a medium size canard and a reduced horizontal tail. The impact of these proposed changes did not affect the TCA-2 test, because it was specifically planned to address power effects on the empennage and a smaller horizontal tail was in the plan to be tested. However, the focus of future tests was reevaluated and the emphasis was shifted away from assessment of TCA specific configurations to a more general assessment of configurations that encompass the projected design space for the TC.

Glessner, Paul T.↗
NASA NTRS1921

Airplane Balance

The authors argue that the center of gravity has a preponderating influence on the longitudinal stability of an airplane in flight, but that manufacturers, although aware of this influence, are still content to apply empirical rules to the balancing of their airplanes instead of conducting wind tunnel tests. The author examines the following points: 1) longitudinal stability, in flight, of a glider with coinciding centers; 2) the influence exercised on the stability of flight by the position of the axis of thrust with respect to the center of gravity and the whole of the glider; 3) the stability on the ground before taking off, and the influence of the position of the landing gear. 4) the influence of the elements of the glider on the balance, the possibility of sometimes correcting defective balance, and the valuable information given on this point by wind tunnel tests; 5) and a brief examination of the equilibrium of power in horizontal flight, where the conditions of stability peculiar to this kind of flight are added to previously existing conditions of the stability of the glider, and interfere in fixing the safety limits of certain evolutions.

Huguet, L↗
NASA NTRS1987

Potential benefits of magnetic suspension and balance systems

The potential of Magnetic Suspension and Balance Systems (MSBS) to improve conventional wind tunnel testing techniques is discussed. Topics include: elimination of model geometry distortion and support interference to improve the measurement accuracy of aerodynamic coefficients; removal of testing restrictions due to supports; improved dynamic stability data; and stores separation testing. Substantial increases in wind tunnel productivity are anticipated due to the coalescence of these improvements. Specific improvements in testing methods for missiles, helicopters, fighter aircraft, twin fuselage transports and bombers, state separation, water tunnels, and automobiles are also forecast. In a more speculative vein, new wind tunnel test techniques are envisioned as a result of applying MSBS, including free-flight computer trajectories in the test section, pilot-in-the-loop and designer-in-the-loop testing, shipboard missile launch simulation, and optimization of hybrid hypersonic configurations. Also addressed are potential applications of MSBS to such diverse technologies as medical research and practice, industrial robotics, space weaponry, and ore processing in space.

Lawing, Pierce L.↗
NASA NTRS1932

Problems concerning the stability and maneuverability of airplanes

The stability of an airplane can be easily determined by wind-tunnel tests, especially by simple tests with models mounted wind-vane fashion. However, each stability curve plotted by this method is valid only for a certain setting of the corresponding control surface, i.e., it characterizes the stability of the airplane with the control stick in a given position. The problems thus defined are studied from the point of view of longitudinal and transverse stability. Directional stability is not included in this study.

Biche, Jean↗
NASA NTRS1974

Effects of reaction control system jet simulation on the stability and control characteristics of a 0.015 scale space shuttle orbiter model tested in the Langley Research Center unitary plan wind tunnel

An experimental investigation was performed in the Langley Research Center Unitary Plan Wind Tunnel (Test 0A70) to obtain the detailed effects that RCS jet flow interactions with local orbiter flow field have on supersonic stability and control characteristics of the space shuttle orbiter. Six-component force data were obtained through an angle-of-attack range from 15 to 35 degrees at angles of sideslip of 0, +5, and -5 degrees. The test was conducted with yaw jet simulation at free-stream Mach numbers of 2.5 and 4.6, simulating SSV re-entry flight conditions at these Mach numbers. In addition to the basic force measurements, fuselage base pressures and pressures on the non-metric RCS pods were obtained.

Daileda, J. J.↗
NASA NTRS1980

Dynamics stability derivatives of space shuttle orbiter obtained from wind-tunnel and approach and landing flight tests

A comparison was made between ground facility measurements, the aerodynamic design data book values, and the dynamic damping derivatives extracted from the space shuttle orbiter approach and landing flight tests. The comparison covers an angle of attack range from 2 deg to 10 deg at subsonic Mach numbers. The parameters of pitch, yaw, and roll damping, as well as the yawing moment due to rolling velocity and rolling moment due to yawing velocity are compared.

Freeman, D. C., Jr.↗
NASA NTRS1945

Wind-tunnel Tests of the 0.15-scale Powered Model of the Fleetwings XBTK-1 Airplane : Longitudinal Stability and Control

An investigation was made of the static longitudinal stability, and control and stall characteristics of XBTK-1 dive bomber. Results indicate that the longitudinal stability will probably be satisfactory for all contemplated flight conditions at the rear-most CG location with elevator both fixed and free. Power effects were small. Sufficient elevator control will be available to trim in any flight condition above the ground. Increasing the slotted flap deflection above 30 degrees only slightly increased the max. lift coefficient.

Boykin, Rebecca I.↗
NASA NTRS1954

Wind-tunnel Tests of a Model of a Wingless Fin-controlled Missile to Obtain Static Stability and Control Characteristics Through a Range of Mach Numbers from 0.5 to 0.88

An investigation at medium to high subsonic speeds has been conducted in the Langley low-turbulence pressure tunnel to determine the static stability and control characteristics and to measure the fin normal forces and moments for a model of a wingless fin-controlled missile. The data were obtained at Reynolds number of 2.1 x 10(6) based on the missile maximum diameter or 17.7 x 10(6) based on missile length; this Reynolds number was found to be large enough to avoid any large scale effects between the test and the expected flight Reynolds number. With the horizontal-fin deflection limited to a maximum of 6 degrees, longitudinally stable and trimmed flight could not be maintained beyond an angle of attack of 17 degrees for a Mach number of 0.88 and beyond 20 degrees for a Mach number of 0.50 for any center-of-gravity location without the use of some auxiliary stability or control device such as jet vanes. Mach number had no appreciable effect on the center-of-pressure positions and only a slight effect on neutral-point position. There was a shift in neutral-point position of about 1 caliber as the angle of attack was varied through the range for which the neutral point could be determined. Yawing the model to angles of sideslip up to 7 degrees had little effect on the longitudinal stability at angles of attack up to 15 degrees; however, above 15 degrees, the effect of sideslip was destabilizing. With the vertical fins at a plus-or-minus 6 degree roll deflection, the rolling moment caused by yawing the model at high angles of attack could be trimmed out up to angles of sideslip of 6.5 degrees and an angle of attack of 26 degrees for a Mach number of 0.50; this range of sideslip angles was reduced to 3 degrees at a Mach number of 0.88. The data indicated that, at lower angles of attack, the trim range extended to higher angles of sideslip. The total normal-force and hinge-moment coefficients for both horizontal fins were slightly nonlinear with both angle-of-attack and fin deflection. The effect of Mach number was to reduce the slopes of the hinge-moment coefficient with angle of attack and deflection angle. In general, the effort of increasing the sideslip angle was to reduce the values of the fin normal-force and hinge-moment coefficients.

Burrows, Dale L↗
NASA NTRS1989

Experimental Aerodynamic Characteristics of a Joined-wing Research Aircraft Configuration

A wind-tunnel test was conducted at Ames Research Center to measure the aerodynamic characteristics of a joined-wing research aircraft (JWRA). This aircraft was designed to utilize the fuselage and engines of the existing NASA AD-1 aircraft. The JWRA was designed to have removable outer wing panels to represent three different configurations with the interwing joint at different fractions of the wing span. A one-sixth-scale wind-tunnel model of all three configurations of the JWRA was tested in the Ames 12-Foot Pressure Wind Tunnel to measure aerodynamic performance, stability, and control characteristics. The results of these tests are presented. Longitudinal and lateral-directional characteristics were measured over an angle of attack range of -7 to 14 deg and over an angle of sideslip range of -5 to +2.5 deg at a Mach number of 0.35 and a Reynolds number of 2.2x10(6)/ft. Various combinations of deflected control surfaces were tested to measure the effectiveness and impact on stability of several control surface arrangements. In addition, the effects on stall and post-stall aerodynamic characteristics from small leading-edge devices called vortilons were measured. The results of these tests indicate that the JWRA had very good aerodynamic performance and acceptable stability and control throughout its flight envelope. The vortilons produced a profound improvement in the stall and post-stall characteristics with no measurable effects on cruise performance.

Smith, Stephen C.↗
NASA NTRSDate not supplied

Experimental Testing of Advanced Generalized Predictive Control for Stability Augmentation and Vibration Reduction of Tiltrotor Aircraft

Generalized Predictive Control (GPC) is an advanced form of an adaptive control algorithm that uses experimentally acquired data to determine the input-output relationship of complex systems through a process called system identification (system ID). GPC has historically been applied to wind tunnel tests of dynamically-scaled tiltrotor aircraft for stability augmentation and vibration reduction since the complex nature of these dynamic systems does not lend itself well to traditional control theory. Advanced GPC (AGPC) improves upon traditional GPC by enabling self-adaptation as conditions change from those used to acquire the system ID and controller performance would normally erode. The present research expands upon previous analytical development and demonstration of AGPC with experimental demonstration. To support AGPC, this present work also identifies and describes figures of merit that define a good working controller and quantifies the uniqueness of the control inputs and quality of the system ID parameters. The present research demonstrates that AGPC consistently performs better than traditional GPC and can successfully adapt to changing conditions.

Active Controls↗
NASA NTRSDate not supplied

Experimental Testing of Advanced Generalized Predictive Control for Stability Augmentation and Vibration Reduction of Tiltrotor Aircraft

Generalized Predictive Control (GPC) is an advanced form of an adaptive control algorithm that uses experimentally acquired data to determine the input-output relationship of complex systems through a process called system identification (system ID). GPC has historically been applied to wind tunnel tests of dynamically-scaled tiltrotor aircraft for stability augmentation and vibration reduction since the complex nature of these dynamic systems does not lend itself well to traditional control theory. Advanced GPC (AGPC) improves upon traditional GPC by enabling self-adaptation as conditions change from those used to acquire the system ID and controller performance would normally erode. The present research expands upon previous analytical development and demonstration of AGPC with experimental demonstration. To support AGPC, this present work also identifies and describes figures of merit that define a good working controller and quantifies the uniqueness of the control inputs and quality of the system ID parameters. The present research demonstrates that AGPC consistently performs better than traditional GPC and can successfully adapt to changing conditions.

Active Controls↗
NASA NTRS2010

Current Methods Modeling and Simulating Icing Effects on Aircraft Performance, Stability, Control

Icing alters the shape and surface characteristics of aircraft components, which results in altered aerodynamic forces and moments caused by air flow over those iced components. The typical effects of icing are increased drag, reduced stall angle of attack, and reduced maximum lift. In addition to the performance changes, icing can also affect control surface effectiveness, hinge moments, and damping. These effects result in altered aircraft stability and control and flying qualities. Over the past 80 years, methods have been developed to understand how icing affects performance, stability, and control. Emphasis has been on wind-tunnel testing of two-dimensional subscale airfoils with various ice shapes to understand their effect on the flowfield and ultimately the aerodynamics. This research has led to wind-tunnel testing of subscale complete aircraft models to identify the integrated effects of icing on the aircraft system in terms of performance, stability, and control. Data sets of this nature enable pilot-in-the-loop simulations to be performed for pilot training or engineering evaluation of system failure impacts or control system design.

Ratvasky, Thomas P.↗
NASA NTRS1981

A full-scale wind tunnel investigation of a helicopter bearingless main rotor

A helicopter bearingless main rotor was tested. Areas of investigation included aeroelastic stability, aerodynamic performance, and rotor loads as a function of collective pitch setting, RPM, airspeed and shaft angle. The rotor/support system was tested with the wind tunnel balance dampers installed and, subsequently, removed. Modifications to the rotor hub were tested. These included a reduction in the rotor control system stiffness and increased flexbeam structural damping. The primary objective of the test was to determine aeroelastic stability of the fundamental flexbeam/blade chordwise bending mode. The rotor was stable for all conditions. Damping of the rotor chordwise bending mode increases with increased collective pitch angle at constant operating conditions. No significant decrease in rotor damping occured due to frequency coalescence between the blade chordwise fundamental bending mode and the support system.

Warmbrodt, W.↗
NASA NTRS2013

Analysis of a Split-Plot Experimental Design Applied to a Low-Speed Wind Tunnel Investigation

A procedure to analyze a split-plot experimental design featuring two input factors, two levels of randomization, and two error structures in a low-speed wind tunnel investigation of a small-scale model of a fighter airplane configuration is described in this report. Standard commercially-available statistical software was used to analyze the test results obtained in a randomization-restricted environment often encountered in wind tunnel testing. The input factors were differential horizontal stabilizer incidence and the angle of attack. The response variables were the aerodynamic coefficients of lift, drag, and pitching moment. Using split-plot terminology, the whole plot, or difficult-to-change, factor was the differential horizontal stabilizer incidence, and the subplot, or easy-to-change, factor was the angle of attack. The whole plot and subplot factors were both tested at three levels. Degrees of freedom for the whole plot error were provided by replication in the form of three blocks, or replicates, which were intended to simulate three consecutive days of wind tunnel facility operation. The analysis was conducted in three stages, which yielded the estimated mean squares, multiple regression function coefficients, and corresponding tests of significance for all individual terms at the whole plot and subplot levels for the three aerodynamic response variables. The estimated regression functions included main effects and two-factor interaction for the lift coefficient, main effects, two-factor interaction, and quadratic effects for the drag coefficient, and only main effects for the pitching moment coefficient.

Erickson, Gary E.↗