Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “angle control”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 1,333 records · Page 74
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

Icing effects on aircraft stability and control determined from flight data - Preliminary results

The effects of airframe icing on the stability and control characteristics of the NASA DH-6 Twin Otter icing research aircraft were investigated by flight test. The flight program was developed to obtain stability and control parameters of the DH-6 in a baseline ('uniced') configuration and an 'artificially iced' configuration for specified thrust conditions. Stability and control parameter identification maneuvers were performed over a wide range of angles of attack for wing flaps retraced (0 deg) and wing flaps partially deflected (10 deg). Engine power was adjusted to hold thrust constant at one of three thrust coefficients (C(sub T) = 0.14, C(sub T) = 0.07, C(sub T) = 0.00). This paper presents only the pitching- and yawing-moment results from the flight test program. Stability and control parameters were estimated for the uniced and artificially iced configurations using a modified stepwise regression algorithm. Comparisons of the uniced and iced stability and control parameters are presented for the majority of the flight envelope. The artificial ice reduced the elevator and rudder control effectiveness by 12 percent and 8 percent respectively for the 0 deg flap setting. The longitudinal static stability was also decreased substantially (approximately 10 percent) because of the tail ice. Further discussion is provided to explain some of the effects of ice on the stability and control parameters.

Ratvasky, T. P.↗
NASA NTRS2002

Ground and Flight Evaluation of a Small-Scale Inflatable-Winged Aircraft

A small-scale, instrumented research aircraft was flown to investigate the night characteristics of innersole wings. Ground tests measured the static structural characteristics of the wing at different inflation pressures, and these results compared favorably with analytical predictions. A research-quality instrumentation system was assembled, largely from commercial off-the-shelf components, and installed in the aircraft. Initial flight operations were conducted with a conventional rigid wing having the same dimensions as the inflatable wing. Subsequent flights were conducted with the inflatable wing. Research maneuvers were executed to identify the trim, aerodynamic performance, and longitudinal stability and control characteristics of the vehicle in its different wing configurations. For the angle-of-attack range spanned in this flight program, measured flight data demonstrated that the rigid wing was an effective simulator of the lift-generating capability of the inflatable wing. In-flight inflation of the wing was demonstrated in three flight operations, and measured flight data illustrated the dynamic characteristics during wing inflation and transition to controlled lifting flight. Wing inflation was rapid and the vehicle dynamics during inflation and transition were benign. The resulting angles of attack and of sideslip ere small, and the dynamic response was limited to roll and heave motions.

Murray, James E.↗
NASA NTRS1984

Investigation of the effect of vehicle, angle-of-attack, and trim elevon position on lateral-directional aerodynamic parameters of the Shuttle Orbiter

This paper discusses the lateral-directional stability and control coefficients extracted from onboard measurements made during reentry of flights two to eight of the Space Shuttle Orbiters. Maximum likelihood estimation is applied to data derived from accelerometer and rate gyro measurements and trajectory, meteorological and control surface data to estimate these lateral-directional parameters. Comparison of estimated coefficients across the seven flights and two shuttle vehicles is made. The effect of the angle-of-attack, and trim elevon deflection on estimated values of the stability and control derivatives is also studied.

Schiess, J. R.↗
NASA NTRS1935

The Aerodynamic Forces and Moments on a Spinning Model of the F4B-2 Airplane as Measured by the Spinning Balance

The aerodynamic forces and moments on a 1/12-scale model of the F4B-2 airplane were measured with the spinning balance in nine spinning attitudes with three sets of tail surfaces, namely, F4B-2 surfaces; F4B-4 fin and rudder with rectangular stabilizer; and with all tail surfaces removed. In one of these attitudes measurements were made to determine the effect upon the forces and moments of independent and of simultaneous displacement of the rudder and elevator for two of the sets of tail surfaces. Additional measurements were made for a comparison of model and full-scale data for six attitudes that were determined from flight tests with various control settings. The characteristics were found to vary in the usual manner with angle of attack and sideslip. The F4B-2 surfaces were quite ineffective as a source of yawing moments. The F4B-4 fin and F4B-2 stabilizer gave a greater damping yawing moment when controls were against the spin than did the F4B-2 surfaces but otherwise there was little difference. Substitution of a rectangular stabilizer for the F4B-2 stabilizer made no appreciable difference in the coefficient. Further comparisons with other airplane types are necessary before final conclusions can be drawn as to the relations between model and full-scale spin measurements.

SPINNING↗
NASA NTRS1980

Rotary balance data for a single-engine agricultural airplane configuration for an angle-of-attack range of 8 deg to 90 deg

Aerodynamic characteristics obtained in a helical flow environment utilizing a rotary balance located in the Langley spin tunnel are presented in plotted form for a 1/10 scale single engine agricultural airplane model. The configurations tested include the basic airplane, various wing leading edge and wing tip devices, elevator, aileron, and rudder control settings, and other modifications. Data are presented without analysis for an angle of attack range of 8 deg to 90 deg, and clockwise and counter-clockwise rotations covering a spin coefficient range from 0 to .9.

Mulcay, W. J.↗
NASA NTRS1993

Vortex developments over steady and accelerated airfoils incorporating a trailing edge jet

Computational and experimental studies are conducted to investigate the influence of a trailing edge jet on flow separation and subsequent vortex formation over steady and accelerated airfoils at high angles of attack. A computer code, employing the stream function-vorticity approach, is developed and utilized to conduct numerical experiments on the flow problem. To verify and economize such efforts, an experimental system is developed and incorporated into a subsonic wind tunnel where streamline and vortex flow visualization experiments are conducted. The study demonstrates the role of the trailing edge jet in controlling flow separation and subsequent vortex development for steady and accelerating flow at angles past the static stall angle of attack. The results suggest that the concept of the trailing edge jet may be utilized to control the characteristics of unsteady separated flows over lifting surfaces. This control possibility seems to be quite effective and could have a significant role in controlling unsteady separated flows.

Finaish, F.↗
NASA NTRS1993

Robust control of hypersonic vehicles considering propulsive and aeroelastic effects

The influence of propulsion system variations and elastic fuselage behavior on the flight control system of an airbreathing hypersonic vehicle is investigated. Thrust vector magnitude and direction changes due to angle of attack variations affect the pitching moment. Low structural vibration frequencies may occur close to the rigid body modes influencing the angle of attack and lead to possible cross coupling. These effects are modeled as uncertainties in the context of a robust control study of a hypersonic vehicle model accelerating through Mach 8 using H-infinity and mu synthesis techniques. Various levels of uncertainty are introduced into the system. Both individual and simultaneous appearance of uncertainty are considered. The results indicate that the chosen design technique is suitable for this kind of problem provided that a fairly good knowledge of the effects mentioned above is available. The order of the designed controller is reduced but robust performance is lost which shows the need for fixed order design techniques.

Buschek, Harald↗
NASA NTRS1992

A satellite anemometer

We present the conceptual design of an instrument that can provide a continuous analog measure of the pitch and yaw angles of a low altitude satellite in the frame of the neutral atmosphere. The device, which uses pressure sensor orientation to provide its signal, can function in an attitude control loop and/or be used to measure atmospheric winds or wave motions. Any arbitrary angle of attack (pitch and yaw) less than +/- 45 deg can be selected for a heading. The sensitivity of the device is 0.1 deg but this would be limited by low ambient pressure at high altitudes and by multiple particle collisions in and below the slip flow region. The development of a prototype flight unit is in progress.

Hanson, W. B.↗
NASA NTRSDate not supplied

Mars Sample Return, Sample Retrieval Lander, Reaction Control System Jet Interaction Supersonic Wind Tunnel Test Overview with CFD Predictions

NASA's Mars Sample Return campaign will be launching several missions over the next decade that will work together to return rock samples from Mars back to Earth. The Sample Retrieval Lander (SRL) will deliver the Mars Ascent Vehicle and fetch rover to the surface of Mars in 2006. Rock samples collected by the Mars 2020 Perseverance rover, landing in early 2021, will be loaded on the the ascent vehicle to be launched into Mars orbit for retrieval by yet another spacecraft. The Sample Retrieval Lander will be a blunt entry capsule similar to past Mars entry vehicles like Mars Science Laboratory and Viking. The vehicle will fly a guided entry, using a small lift vector produced by a non-zero trim angle of attack to eliminate downrange and crossrange position errors at the point of parachute deploy. This energy and heading management is achieved with a reaction control system (RCS) that directs the bank angle of the vehicle and also minimizes unwanted capsule dynamics. The reaction control system and control design is based on the Mars Science Laboratory and Mars 2020 RCS systems. However, due to packaging constraints, the backshell of this new entry vehicle has a different geometry than those earlier designs. To certify the RCS system for flight the project must characterize the jet plume interactions with the capsule backshell that could impair or significantly augment the RCS control authority. This characterization will be done through a combination of computational fluid dynamics (CFD) analysis and wind tunnel test. Two candidate arrangements of the RCS jets have been identified for the SRL vehicle and are currently under evaluation before final selection. The aero/RCS plume interactions of these candidate configurations have been measured in a supersonic wind tunnel test in NASA Langley's Unitary Plan Wind Tunnel. The test was conducted in the fall of 2020 and data is currently being reduced. An overview of the candidate RCS configurations are presented here with an overview of the wind tunnel model design, jet scaling and scaled nozzle design, and the test matrix. Preliminary CFD runs are presented with an assessment of the predicted plumes and their interaction with the wake flow of the vehicle. The predicted effects of the model sting is provided as well. This high fidelity wind tunnel test is being conducted much earlier in the SRL project than would normally be done. The test was funded as part of a CFD evaluation task funded by NASA's Aerosciences Evaluation and Test Capabilities Project. The objective of the evaluation task was to compare the ability of CFD to predict complex flows with data that can be measured in the Langley Unitary Plan Wind Tunnel. RCS Jet interactions were selected as a type of complex flow that is important to NASA missions. In addition to providing useful data to the SRL project, there was added emphasis on quantifying the accuracy of the CFD predictions and wind tunnel test data. An overview of the uncertainty quantification methodologies for computational and experimental portions of this test is presented.

blunt body↗
NASA NTRS1992

Dynamic Forms and Their Application to Control

The rigid body model plays a basic role in both mod- and control of many systems, including aircraft flight control. The distinctive feature of the rigid body is that its motion evolves on a curved space, Consequently, the designer must take into account the deterioration of chosen coordinations near their singularities. In these regions the control actions and the sensitivity to modeling and sensor errors become unacceptable. Thus, coordinate patches and patching logic should be a part of the overall control law design and implementation. The various dynamic forms, such as scalar forms, vector forms, matrix forms, rotational forms, Euler angle forms, and Euler parameter forms, described in the paper provide an effective means for coordinate patching and control law implementation. The utility of these forms is illustrated by means of several applications to flight control.

Meyer, George↗
NASA NTRS1946

High-Speed Longitudinal-Stability and Control Characteristics of the Consolidated Vultee Lark Missile as Predicted from Wind-Tunnel Tests (TED No. NACA 2391)

A high-speed wind-tunnel investigation of the aerodynamic characteristics of a full-scale model of the Consolidated Vultee Lark indicates that the missile possesses satisfactory longitudinal-stability and-control characteristics throughout the Mach number range from 0.2 to 0.85, but that the maximum lift coefficients developed are not high enough to insure interception of the target at high altitudes. A reduction in wing loading appears advisable. Although the static longitudinal stability at zero angle of attack changes with Mach number and with lift coefficient, satisfactory control should be possible at all times as the tails retain their relatively large effectiveness throughout the range of Mach numbers and lift coefficients tested. Minimum stability and maximum maneuverability occur around 0.80 Mach number and 0.2 lift coefficient, which corresponds to level flight conditions of the missile. The optimum ratio of tail-to-wing deflection is 0.4.

Axelson, John A.↗
NASA NTRS2018

Robust Bonding Through Process Control

Process control has been proven to be the most reliable means of safeguarding the quality of adhesive bonds according to Federal Aviation Administration (FAA). A method for implementing process control for reduction in risk in a bonded joint fabrication process is demonstrated in this study using a selected bonding system. The stepwise method included risk analysis to identify defects with the highest impact and likelihood to occur, evaluation of various pre-bond surface analysis tools to monitor for the selected defects, and demonstration of the benefits of in-process monitoring utilizing threshold limits determined from bond performance tests. The bonded system selected for investigation was an aerospace carbon fiber epoxy composite substrate surface prepared with random orbital sanding using 180 grit aluminum oxide sand paper. A series of portable, pre-bond surface analysis tools were investigated for their ability to be used for in-line bond process control. Results and threshold limits are presented from roughness, ballistic water contact angle (WCA), color, gloss, and Fourier transform infrared spectroscopy (FTIR) surface analysis tools. Results demonstrated how in-process inspection methods can be used to ensure quality of a surface preparation for a selected bonding system. A framework is provided for implementation of bond process control for robust bonding.

Kutscha, Eileen O.↗
NASA NTRS1985

Static internal performance of a single-engine onaxisymmetric-nozzle vaned-thrust-reverser design with thrust modulation capabilities

An investigation has been conducted at wind-off conditions in the stati-test facility of the Langley 16-Foot Transonic Tunnel. The tests were conducted on a single-engine reverser configuration with partial and full reverse-thrust modulation capabilities. The reverser design had four ports with equal areas. These ports were angled outboard 30 deg from the vertical impart of a splay angle to the reverse exhaust flow. This splaying of reverser flow was intended to prevent impingement of exhaust flow on empennage surfaces and to help avoid inlet reingestion of exhaust gas when the reverser is integrated into an actual airplane configuration. External vane boxes were located directly over each of the four ports to provide variation of reverser efflux angle from 140 deg to 26 deg (measured forward from the horizontal reference axis). The reverser model was tested with both a butterfly-type inner door and an internal slider door to provide area control for each individual port. In addition, main nozzle throat area and vector angle were varied to examine various methods of modulating thrust levels. Other model variables included vane box configuration (four or six vanes per box), orientation of external vane boxes with respect to internal port walls (splay angle shims), and vane box sideplates. Nozzle pressure ratio was varied from 2.0 approximately 7.0.

Leavitt, L. D.↗
NASA NTRS1979

Vortex-lift roll-control device

A wing is described for aircraft of cropped, arrow-type planform with thin leading and side edges. The wing has a pivotable tip to alter the crop angle of the wing during flight. Increasing the crop angle causes the wing side edge to become a trailing edge which reduces the strength of the side edge vortex flow. Decreasing the crop angle causes opposite results, in particular the side edge is now a leading edge and can generate a leading edge vortex flow. The wing constitutes a roll control device for aircraft of the stated design particularly effective at higher angles of attack.

Lamar, J. E.↗
NASA NTRS1987

Feedback control for penetrating a downburst

Feedback logic is given that produces satisfactory penetration of an aircraft through a severe downburst. The logic has two parts: (1) A switch that calls for full throttle and tight climb-rate-hold feedback logic when the magnitude of headwind-rate exceeds a threshold level; this switch stays on until re-set by the pilot after penetration of the downburst. (2) A tight climb-rate-hold autopilot, using feedback of pitch-angle to angle-of-attack. After penetration, recovery to steady climbing flight is aided by using feedback of airspeed deviation to throttle. Insights on control strategy are gained from an analysis of energy changes during flight through a downburst. Closed-loop simulations are given for a 727 aircraft penetrating severe downbursts on take-off. Boeing's recommendation of constant pitch angle was confirmed as an excellent feedback control strategy.

Bryson, Arthur E., Jr.↗
NASA NTRS1978

Space shuttle orbiter reaction control system interactions with the vehicle flow field

Test data on the effect of the interaction between the space shuttle orbiter rear-mounted reaction control system jets and the flow over the vehicle in the atmosphere are analyzed with particular attention to yaw angle effects. Masted tank-plus-orbiter jet interactions resulting from combinations of forward and rear-mounted reaction control jets being fired prior to external tank staging are also documented. The primary correlating parameter used for the mated data is a single-jet momentum ratioed to free-stream momentum. An analytic model for computing reaction control system interactions for all control combinations used for mated vehicle coast is included.

Rausch, J. R.↗
NASA NTRS1991

Buffet induced structural/flight-control system interaction of the X-29A aircraft

High angle-of-attack flight regime research is currently being conducted for modern fighter aircraft at the NASA Ames Research Center's Dryden Flight Research Facility. This flight regime provides enhanced maneuverability to fighter pilots in combat situations. Flight research data are being acquired to compare and validate advanced computational fluid dynamic solutions and wind-tunnel models. High angle-of-attack flight creates unique aerodynamic phenomena including wing rock and buffet on the airframe. These phenomena increase the level of excitation of the structural modes, especially on the vertical and horizontal stabilizers. With high gain digital flight-control systems, this structural response may result in an aeroservoelastic interaction. A structural interaction on the X-29A aircraft was observed during high angle-of-attack flight testing. The roll and yaw rate gyros sensed the aircraft's structural modes at 11, 13, and 16 Hz. The rate gyro output signals were then amplified through the flight-control laws and sent as commands to the flaperons and rudder. The flight data indicated that as the angle of attack increased, the amplitude of the buffet on the vertical stabilizer increased, which resulted in more excitation to the structural modes. The flight-control system sensors and command signals showed this increase in modal power at the structural frequencies up to a 30 degree angle-of-attack. Beyond a 30 degree angle-of-attack, the vertical stabilizer response, the feedback sensor amplitude, and control surface command signal amplitude remained relatively constant. Data are presented that show the increased modal power in the aircraft structural accelerometers, the feedback sensors, and the command signals as a function of angle of attack. This structural interaction is traced from the aerodynamic buffet to the flight-control surfaces.

Voracek, David F.↗