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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1996

Numerical Simulation of Forced and Free-to-Roll Delta-Wing Motions

The three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate nonsteady vortical flow about a 65-deg sweep delta wing at 30-deg angle of attack. Two large-amplitude, high-rate, forced-roll motions, and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are in good agreement with the forces, moments, and roll-angle time histories. Vortex breakdown is present in each case. Significant time lags in the vortex breakdown motions relative to the body motions strongly influence the dynamic forces and moments.

Chaderjian, Neal M.↗
NASA NTRS1959

Low-Subsonic Measurements of the Static and Oscillatory Lateral Stability Derivatives of a Sweptback-Wing Airplane Configuration at Angles of Attack from -10 to 90 deg

An investigation has been conducted in the Langley free-flight tunnel at low-subsonic speeds to provide some basic information on the stability and control characteristics in the high angle-of-attack range of an airplane configuration typical of current design trends. The investigation consisted of static- and dynamic-force tests over an angle-of- attack range from -10 to 90 deg. The dynamic-force tests, which consisted of both linear- and rotary-oscillation tests, were conducted at values of the reduced-frequency parameter k of 0.10, 0.15, and 0.20. The configuration was directionally unstable for all angles of attack above about 15 deg but maintained positive effective dihedral, control effectiveness, and damping in roll and yaw over most of the angle-of-attack range tested. The effects of frequency on the oscillatory stability derivatives were found to be generally small, but in a few cases the effects were relatively large.

Hewes, Donald E.↗
NASA NTRS1959

Effect of Artificial Pitch Damping on the Longitudinal and Rolling Stability of Aircraft with Negative Static Margins

A preliminary theoretical investigation has been made of the short-period longitudinal and steady-rolling (inertia coupling) stability of a hypersonic glider configuration for center-of-gravity locations rear-ward of the airplane neutral point. Such center-of-gravity positions for subsonic flight would improve performance by reducing supersonic and hypersonic static margins and trim drag. Results are presented of stability calculations and a simulator study for a velocity of 700 ft/sec and an altitude of 401,000 feet. With no augmentation, the airplane was rapidly divergent and was considered unsatisfactory in the simulator study. When a pitch damper was employed as a stability augmenter, the short-period mode became overdamped, and the airplane was easily controlled on the simulator. A steady-rolling analysis showed that the airplane can be made free of rolling divergence for all roll rates with an appropriate damper gain.

Moul, Martin T.↗
NASA NTRS1959

Predicted Static Aeroelastic Effects on Wings with Supersonic Leading Edges and Streamwise Tips

A method is presented for calculation of static aeroelastic effects on wings with supersonic leading edges and streamwise tips. Both chord-wise and spanwise deflections are taken into account. Aerodynamic and structural forces are introduced in influence coefficient form; the former are developed from linearized supersonic wing theory and the latter are assumed to be known from load-deflection tests or theory. The predicted effects of flexibility on lateral-control effectiveness, damping in roll, and lift-curve slope are shown for a low-aspect-ratio wing at Mach numbers of 1.25 and 2.60. The control effectiveness is shown for a trailing-edge aileron, a tip aileron, and a slot-deflector spoiler located along the 0.70 chord line. The calculations indicate that the tip aileron is particularly attractive from an aeroelastic standpoint, because the changes in effectiveness with dynamic pressure are small compared to the changes in effectiveness of the trailing-edge aileron and slot-deflector spoiler. The effects of making several simplifying assumptions in the example calculations are shown. The use of a modified strip theory to determine the aerodynamic influence coefficients gave adequate results only for the high Mach number case. Elimination of chordwise bending in the structural influence coefficients exaggerated the aeroelastic effects on rolling-moment and lift coefficients for both Mach numbers.

Brown, Stuart C.↗
NASA NTRS2000

Large-Amplitude, High-Rate Roll Oscillations of a 65 deg Delta Wing at High Incidence

The IAR/WL 65 deg delta wing experimental results provide both detail pressure measurements and a wide range of flow conditions covering from simple attached flow, through fully developed vortex and vortex burst flow, up to fully-stalled flow at very high incidence. Thus, the Computational Unsteady Aerodynamics researchers can use it at different level of validating the corresponding code. In this section a range of CFD results are provided for the 65 deg delta wing at selected flow conditions. The time-dependent, three-dimensional, Reynolds-averaged, Navier-Stokes (RANS) equations are used to numerically simulate the unsteady vertical flow. Two sting angles and two large- amplitude, high-rate, forced-roll motions and a damped free-to-roll motion are presented. The free-to-roll motion is computed by coupling the time-dependent RANS equations to the flight dynamic equation of motion. The computed results are compared with experimental pressures, forces, moments and roll angle time history. In addition, surface and off-surface flow particle streaks are also presented.

Chaderjian, Neal M.↗
NASA NTRS1954

Flight Tests of a 0.13-Scale Model of the Convair XFY-1 Vertically Rising Airplane with the Lower Vertical Tail Removed, TED No.DE 368

An experimental investigation has been conducted to determine the dynamic stability and control characteristics in hovering and transition flight of a 0.13-scale flying model of the Convair XFY-1 vertically rising airplane with the lower vertical tail removed. The purpose of the tests was to obtain a general indication of the behavior of a vertically rising airplane of the same general type as the XFY-1 but without a lower vertical tail in order to simplify power-off belly landings in an emergency. The model was flown satisfactorily in hovering flight and in the transition from hovering to normal unstalled forward flight (angle of attack approximately 30deg). From an angle of attack of about 30 down to the lowest angle of attack covered in the flight tests (approximately 15deg) the model became progressively more difficult to control. These control difficulties were attributed partly to a lightly damped Dutch roll oscillation and partly to the fact that the control deflections required for hovering and transition flight were too great for smooth flight at high speeds. In the low-angle-of-attack range not covered in the flight tests, force tests have indicated very low static directional stability which would probably result in poor flight characteristics. It appears, therefore, that the attainment of satisfactory directional stability, at angles of attack less than 10deg, rather than in the hovering and transition ranges of flight is the critical factor in the design of the vertical tail for such a configuration.

Lovell, Powell M., Jr.↗
NASA NTRS1948

An Investigation of the Wing and the Wing-Fuselage Combination of a Full-Scale Model of the Republic XP-91 Airplane in the Ames 40-by 80-Foot Wing Tunnel

Wind-tunnel tests of a full-scale model of the Republic XP-91 airplane were conducted to determine the longitudinal and lateral characteristics of the wing alone and the wing-fuselage combination, the characteristics of the aileron, and the damping in roll af the wing alone. Various high-lift devices were investigated including trailing-edge split flaps and partial- and full-span leading-edge slats and Krueger-type nose flaps. Results of this investigation showed that a very significant gain in maximum lift could be achieved through use of the proper leading-edge device, The maximum lift coefficient of the model with split flaps and the original partial-span straight slats was only 1.2; whereas a value of approximately 1.8 was obtained by drooping the slat and extending it full span, Improvement in maximum lift of approximately the same amount resulted when a full-span nose flap was substituted for the original partial-span slat.

Hunton, Lynn W.↗
NASA NTRS1944

A Theoretical Investigation of the Rolling Oscillations of an Airplane with Ailerons Free

An analysis is made of the stability of an airplane with ailerons free, with particular attention to the motions when the ailerons have a tendency to float against the wind. The present analysis supersedes the aileron investigation contained in NACA Report No. 709. The equations of motion are first written to include yawing and sideslipping, and it is demonstrated that the principal effects of freeing the ailerons can be determined without regard to these motions. If the ailerons tend to float against the wind and have a high degree of aerodynamic balance, rolling oscillations, in addition to the normal lateral oscillations, are likely to occur. On the basis of the equations including only the rolling motion and the aileron deflection, formulas are derived for the stability and damping of the rolling oscillations in terms of the hinge moment derivatives and other characteristics of the ailerons and airplane. Charts are also presented showing the oscillatory regions and stability boundaries for a fictitious airplane of conventional proportions. The effects of friction in the control system are investigated and discussed. If the ailerons tend to trail with the wind, the condition for stable variation of stick force with aileron deflection is found to determine the amount of aerodynamic balance that may be used. If the ailerons tend to float against the wind, the period and damping of the rolling oscillations are found to be satisfactory (in a mass-balanced system) so long as the restoring moment is not completely balanced out. Unbalanced mass behind the hinge, however, has an unfavorable effect on the damping of the oscillations and so shifts the boundary that close aerodynamic balance may not be attainable.

Cohen, Doris↗
NASA NTRS1975

The effects of stability augmentation on the gust response of a STOL aircraft during a curved manual approach

The task of investigating the effect of SAS was formulated as two optimal control problems for stochastic systems: (1) to compute SAS gains with a rate model in the performance index algorithm; and (2) to calculate the pilot gains and system gust response using an optimal pilot model. Both problems were solved to yield reasonable low gains for the pilot and SAS, and the lateral directional mode poles and the longitudinal short period poles could be placed accurately by the model matching algorithm. Of the longitudinal (SAS) poles achieved, the vertical rms path error was least for the unaugmented poles. The lateral rms path error was an order of magnitude larger than the vertical error and showed a plus or minus 50 percent variation with SAS. It increased with dutch roll frequency and damping, and it decreased most significantly with increased roll stability. The variation in lateral error with bank angle for curved flight was also a significant function of the augmented poles.

Porter, M. B., Jr.↗
NASA NTRS1981

Experimental and analytical studies of advanced air cushion landing systems

Several concepts are developed for air cushion landing systems (ACLS) which have the potential for improving performance characteristics (roll stiffness, heave damping, and trunk flutter), and reducing fabrication cost and complexity. After an initial screening, the following five concepts were evaluated in detail: damped trunk, filled trunk, compartmented trunk, segmented trunk, and roll feedback control. The evaluation was based on tests performed on scale models. An ACLS dynamic simulation developed earlier is updated so that it can be used to predict the performance of full-scale ACLS incorporating these refinements. The simulation was validated through scale-model tests. A full-scale ACLS based on the segmented trunk concept was fabricated and installed on the NASA ACLS test vehicle, where it is used to support advanced system development. A geometrically-scaled model (one third full scale) of the NASA test vehicle was fabricated and tested. This model, evaluated by means of a series of static and dynamic tests, is used to investigate scaling relationships between reduced and full-scale models. The analytical model developed earlier is applied to simulate both the one third scale and the full scale response.

Lee, E. G. S.↗
NASA NTRS1944

A theoretical investigation of the rolling oscillations of an airplane with ailerons free

An analysis is made of the stability of an airplane with ailerons free, with particular attention to the motions when the ailerons have a tendency to float against the wind. The present analysis supersedes the aileron investigation contained in NACA Technical Report no. 709. The equations of motion are first written to include yawing and sideslipping, and it is demonstrated that the principal effects of freeing the ailerons can be determined without regard to these motions. If the ailerons tend to float against the wind and have a high degree of aerodynamic balance, rolling oscillations, in addition to the normal lateral oscillations, are likely to occur. On the basis of the equations including only the rolling motion and the aileron deflection, formulas derived for the stability and damping of the rolling oscillations in terms of the hinge-moment derivatives are also presented showing the oscillatory regions and stability boundaries for a fictitious airplane of conventional proportion. The effects of friction in the control system are investigated and discussed.

Cohen, Doris↗
NASA NTRS1961

Investigation of Low-Subsonic Flight Characteristics of a Model of a Hypersonic Boost-Glide Configuration Having a 78 deg. Delta Wing

An investigation of the low-subsonic stability and control characteristics of a model of a hypersonic boost-glide configuration having 78 deg. sweep of the leading edge has been made in the Langley full-scale tunnel. The model was flown over an angle-of-attack range from 10 to 35 deg. Static and dynamic force tests were made in the Langley free-flight tunnel. The investigation showed that the longitudinal stability and control characteristics were generally satisfactory with neutral or positive static longitudinal stability. The addition of artificial pitch damping resulted in satisfactory longitudinal characteristics being obtained with large amounts of static instability. The most rearward center-of-gravity position for which sustained flights could be made either with or without pitch damper corresponded to the calculated maneuver point. The lateral stability and control characteristics were satisfactory up to about 15 deg. angle of attack. The damping of the Dutch roll oscillation decreased with increasing angle of attack; the oscillation was about neutrally stable at 20 deg. angle of attack and unstable at angles of attack of about 25 deg. and above. Artificial damping in roll greatly improved the lateral characteristics and resulted in flights being made up to 35 deg. angle of attack.

Paulson, John W.↗
NASA NTRS1959

Investigation of Low-Subsonic Flight Characteristics of a Model of a Flat-Bottom Hypersonic Boost-Glide Configuration Having a 78 deg Delta Wing

An investigation of the low-subsonic stability and control characteristics of a model of a flat-bottom hypersonic boost-glide configuration having 78 deg sweep of the leading edge has been made in the Langley full-scale tunnel. The model was flown over an angle-of-attack range from 10 to 35 deg. Static and dynamic force tests were made in the Langley free-flight tunnel. The investigation showed that the longitudinal stability and control characteristics were generally satisfactory with neutral or positive static longitudinal stability. The addition of artificial pitch damping resulted in satisfactory longitudinal characteristics being obtained with large amounts of static instability. The most rearward center-of-gravity position for which sustained flights could be made either with or without pitch damper corresponded to the calculated maneuver point. The lateral stability and control characteristics were satisfactory up to about 15 deg angle of attack. The damping of the Dutch roll oscillation decreased with increasing angle of attack; the oscillation was about neutrally stable at 20 deg angle of attack and unstable at angles of attack of about 25 deg and above. Artificial damping in roll greatly improved the lateral characteristics and resulted in flights being made up to 35 deg angle of attack.

Paulson, John W.↗
NASA NTRS1993

Conical Euler analysis and active roll suppression for unsteady vortical flows about rolling delta wings

A conical Euler code was developed to study unsteady vortex-dominated flows about rolling, highly swept delta wings undergoing either forced motions or free-to-roll motions that include active roll suppression. The flow solver of the code involves a multistage, Runge-Kutta time-stepping scheme that uses a cell-centered, finite-volume, spatial discretization of the Euler equations on an unstructured grid of triangles. The code allows for the additional analysis of the free to-roll case by simultaneously integrating in time the rigid-body equation of motion with the governing flow equations. Results are presented for a delta wing with a 75 deg swept, sharp leading edge at a free-stream Mach number of 1.2 and at 10 deg, 20 deg, and 30 deg angle of attack alpha. At the lower angles of attack (10 and 20 deg), forced-harmonic analyses indicate that the rolling-moment coefficients provide a positive damping, which is verified by free-to-roll calculations. In contrast, at the higher angle of attack (30 deg), a forced-harmonic analysis indicates that the rolling-moment coefficient provides negative damping at the small roll amplitudes. A free-to-roll calculation for this case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation, which is characteristic of highly swept delta wings. This limit cycle oscillation may be actively suppressed through the use of a rate-feedback control law and antisymmetrically deflected leading-edge flaps. Descriptions of the conical Euler flow solver and the free-to roll analysis are included in this report. Results are presented that demonstrate how the systematic analysis of the forced response of the delta wing can be used to predict the stable, neutrally stable, and unstable free response of the delta wing. These results also give insight into the flow physics associated with unsteady vortical flows about delta wings undergoing forced motions and free-to-roll motions, including the active suppression of the wing-rock type phenomenon. The conical Euler methodology developed is directly extend able to three-dimensional calculations.

Lee-Rausch, Elizabeth M.↗
NASA NTRS1991

Conical Euler methodology for unsteady vortical flows about rolling delta wings

A conical Euler methodology was developed to study unsteady vortex-dominated flows about rolling highly swept delta wings undergoing either forced or free-to-roll motions. The flow solver of the code involves a multistage Runge-Kutta time-stepping scheme which uses a finite-volume spatial discretization of the Euler equations on an unstructured grid of triangles. Results are presented for a 75-deg swept sharp-leading-edge delta wing at a freestream Mach number of 1.2 and at alpha = 10, 20, and 30 deg. At the 10 and 20 deg, forced harmonic analyses indicate that the rolling moment coefficients provide a positive damping which is verified by free-to-roll calculations. In contrast, at 30 deg, a forced harmonic analysis indicates that the rolling moment coefficient provides a negative damping at the small roll amplitudes. A free-to-roll calculation for this case produces an initially divergent response, but as the amplitude of motion grows with time, the response transitions to a wing-rock type of limit cycle oscillation.

Lee, Elizabeth M.↗
NASA NTRS1976

Supersonic dynamic-stability derivatives of the space shuttle launch vehicle

An investigation was made to determine the dynamic stability characteristics of a 0.015-scale model of the space shuttle launch vehicle at supersonic speeds. These tests were made at angles of attack from -10 deg to 10 deg at Mach numbers of 2.00, 2.86, 3.96, and 4.63. The complete launch vehicle, consisting of the orbiter, external tank, and solid rocket boosters, has positive damping in pitch, roll, and yaw over the angle-of-attack and Mach number range. The orbiter external tank configuration has a region of negative angles pitch damping for small negative of attack at a Mach number of 2.00. At all other test conditions, the orbiter external tank configuration has positive damping about all three axes.

Boyden, R. P.↗