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At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS1995

SOFIA 2 model telescope wind tunnel test report

This document outlines the tests performed to make aerodynamic force and torque measurements on the SOFIA wind tunnel model telescope. These tests were performed during the SOFIA 2 wind tunnel test in the 14 ft wind tunnel during the months of June through August 1994. The test was designed to measure the dynamic cross elevation moment acting on the SOFIA model telescope due to aerodynamic loading. The measurements were taken with the telescope mounted in an open cavity in the tail section of the SOFIA model 747. The purpose of the test was to obtain an estimate of the full scale aerodynamic disturbance spectrum, by scaling up the wind tunnel results (taking into account differences in sail area, air density, cavity dimension, etc.). An estimate of the full scale cross elevation moment spectrum was needed to help determine the impact this disturbance would have on the telescope positioning system requirements. A model of the telescope structure, made of a light weight composite material, was mounted in the open cavity of the SOFIA wind tunnel model. This model was mounted via a force balance to the cavity bulkhead. Despite efforts to use a 'stiff' balance, and a lightweight model, the balance/telescope system had a very low resonant frequency (37 Hz) compared to the desired measurement bandwidth (1000 Hz). Due to this mechanical resonance of the balance/telescope system, the balance alone could not provide an accurate measure of applied aerodynamic force at the high frequencies desired. A method of measurement was developed that incorporated accelerometers in addition to the balance signal, to calculate the aerodynamic force.

Keas, Paul↗
NASA NTRS1924

Tests on a model of the D.F.W. airplane T-29 of the "Deutsche Flugzeug derke" (German airplane works)

Experiments similar to those carried out with the A.E.G (Allgemeine Elektrizitats-Gesellschaft) were made in the small wind tunnel of the Gottingen laboratory on a model of the D.F.W. airplane T-29. Three series of tests were carried out on the model with a velocity head (or dynamic pressure) of 5 kg/sq m (1.02 lb/sq ft), during which one of the movable surfaces was deflected at various angles, while both the others were retained in their central positions. Of special interest among the results of the tests is the different run of the elevating moments. The curves for the A.E.G. model, rising to the right, denote stability with the elevator locked, while the slight inclination to the left with the D.F.W model denotes a slight instability. For the maximum C(sub L) values, the stability of A.E.G. model continues to increase and the instability of the D.F.W. model is converted into stability. The rolling moments shown when the angular deflection of the ailerons is 0 degrees are due, in both series of tests, to the unequal distribution of the air velocity over the cross section of the wind tunnel, rather than to a lack of symmetry in the model.

Molthan, Wilhelm↗
NASA NTRS1927

Experiments with three horizontal empennages

The main objective of these experiments was to determine how the air forces on the whole empennage are affected by varying the size of the elevator without changing the size of the whole empennage. The secondary objective was to determine how the magnitude of the air forces acting on the elevator (or the elevator moments) are affected by varying the angle of attack of the whole empennage, or of the elevator alone.

CONTROL UNITS - WIND TUNNEL TESTS↗
NASA NTRS1941

Analysis and prediction of longitudinal stability of airplanes

An analysis has been made of the longitudinal stability characteristics of 15 airplanes as determined in flight. In the correlation of satisfactory and unsatisfactory characteristics with determined values, the derivative that expresses the ratio of static-restoring moments to elevator-control moments was found to represent most nearly the stability characteristics appreciated by the pilots. The analysis was extended to study the effects of various design features on the observed stability characteristics. Design charts and data are included that show the effects on longitudinal stability of relative positions of wing and tail, fuselage size and location, engine nacelles, and horizontal-tail arrangements.

Gilruth, R R↗
NASA NTRS1927

Approximation method for determining the static stability of a monoplane glider

The calculations in this paper afford an approximate solution of the static stability. A derivation of the formulas for moment coefficient of a wing, moment coefficient of elevator, and the total moment of the combined wing and elevator and the moment coefficient with reference to the center of gravity are provided.

AIRFOILS - GLIDERS↗
NASA NTRS1949

Wind-Tunnel Investigation of the Low-Speed Static Stability and Control Characteristics of a Model of the Bell MX-776 (RASCAL) in Combined Angle of Attack and Sideslip

An investigation has been made in the Langley stability tunnel to determine the low-speed static stability and control characteristics of a model of the Bell MX-776. The results show the model to be longitudinally unstable in the angle-of-attack range around zero angle of attack and to become stable at moderate angles of attack. The results of the present investigation agree reasonably well with results obtained in other facilities at low speed. The present pitching-moment results at low Mach numbers also agree reasonably well with unpublished results of tests of the model at supersonic Mach numbers (up to Mach number 1.86). Unpublished results at moderate and high subsonic speeds, however, indicate considerably greater instability at low angles of attack than is indicated by low-speed results. The results of the present tests also showed that the pitching-moment coefficients for angles of attack up to 12deg remained fairly constant with sideslip angle up to 12deg. The elevators tested produced relatively large pitching moments at zero angle of attack but, as the angle of attack was increased, the elevator effectiveness decreased. The rate of decrease of elevator effectiveness with angle of attack was less for 8deg than for 20deg elevator deflection. Therefore although 8deg deflection caused an appreciable change in longitudinal trim angle and trim lift coefficient a deflection of 20deg caused only a small additional increase in trim angle and trim lift coefficient.

Letko, William↗
NASA NTRS1972

Moment expansion for ionospheric range error

On a plane earth, the ionospheric or tropospheric range error depends only on the total refractivity content or zeroth moment of the refracting layer and the elevation angle. On a spherical earth, however, the dependence is more complex; so for more accurate results it has been necessary to resort to complex ray-tracing calculations. A simple, high-accuracy alternative to the ray-tracing calculation is presented. By appropriate expansion of the angular dependence in the ray-tracing integral in a power series in height, an expression is obtained for the range error in terms of a simple function of elevation angle, E, at the expansion height and of the mth moment of the refractivity, N, distribution about the expansion height. The rapidity of convergence is heavily dependent on the choice of expansion height. For expansion heights in the neighborhood of the centroid of the layer (300-490 km), the expansion to N = 2 (three terms) gives results accurate to about 0.4% at E = 10 deg. As an analytic tool, the expansion affords some insight on the influence of layer shape on range errors in special problems.

Mallinckrodt, A.↗
NASA NTRS1934

The Aerodynamic Forces and Moments Exerted on a Spinning Model of the NY-1 Airplane as Measured by the Spinning Balance

A preliminary investigation of the effects of changes in the elevator and rudder settings and of small changes in attitude upon the aerodynamic forces and moments exerted upon a spinning airplane was undertaken with the spinning balance in the 5-foot vertical tunnel of the National Advisory Committee for Aeronautics. The tests were made on a 1/12-scale model of the ny-1 airplane. Data by which to fix the attitude, the radius of spin, and the rotational and air velocities were taken from recorded spins of the full-scale airplane. Two spinning conditions were investigated. All six components of the aerodynamic reaction were measured and are presented in coefficient form refereed to airplane axes. The results indicate that the change in yawing moment produced by the rudder with the elevator up was the only component of force or moment produced by the elevator and rudder that could not have been balanced in an actual spin by small changes in attitude and angular velocity.

Bamber, M J↗
NASA NTRS1942

The Porpoising Characteristics of a Planing Surface Representing the Forebody of a Flying-Boat Hull

Porpoising characteristics were observed on V-body fitted with tail surfaces for different combinations of load, speed, moment of inertia, location of pivot, elevator setting, and tail area. A critical trim was found which was unaltered by elevator setting or tail area. Critical trim was lowered by moving pivot either forward or down or increasing radius or gyration. Increase in mass and moment of inertia increased amplitude of oscillations. Complete results are tabulated and shown graphically.

HULL BOTTOMS, VEE↗
NASA NTRS1992

A preliminary look at an optimal multivariable design for propulsion-only flight control of jet-transport aircraft

Control of a large jet transport aircraft without the use of conventional control surfaces was studied. Engine commands were used to attempt to recreate the forces and moments typically provided by the elevator, ailerons, and rudder. Necessary conditions for aircraft controllability were developed pertaining to aircraft configuration such as the number of engines and engine placement. An optimal linear quadratic regulator controller was developed for the Boeing 707-720, in particular, for regulation of its natural dynamic modes. The design used a method of assigning relative weights to the natural modes, i.e., phugoid and dutch roll, for a more intuitive selection of the cost function. A prototype pilot command interface was then integrated into the loop based on pseudorate command of both pitch and roll. Closed loop dynamics were evaluated first with a batch linear simulation and then with a real time high fidelity piloted simulation. The NASA research pilots assisted in evaluation of closed loop handling qualities for typical cruise and landing tasks. Recommendations for improvement on this preliminary study of optimal propulsion only flight control are provided.

Azzano, Christopher P.↗
NASA NTRS2002

Optimal Aircraft Control Upset Recovery With and Without Component Failures

This paper treats the problem of recovering sustainable nondescending (safe) flight in a transport aircraft after one or more of its control effectors fail. Such recovery can be a challenging goal for many transport aircraft currently in the operational fleet for two reasons. First, they have very little redundancy in their means of generating control forces and moments. These aircraft have, as primary control surfaces, a single rudder and pairwise elevators and aileron/spoiler units that provide yaw, pitch, and roll moments with sufficient bandwidth to be used in stabilizing and maneuvering the airframe. Beyond this, throttling the engines can provide additional moments, but on a much slower time scale. Other aerodynamic surfaces, such as leading and trailing edge flaps, are only intended to be placed in a position and left, and are, hence, very slow-moving. Because of this, loss of a primary control surface strongly degrades the controllability of the vehicle, particularly when the failed effector becomes stuck in a non-neutral position where it exerts a disturbance moment that must be countered by the remaining operating effectors. The second challenge in recovering safe flight is that these vehicles are not agile, nor can they tolerate large accelerations. This is of special importance when, at the outset of the recovery maneuver, the aircraft is flying toward the ground, as is frequently the case when there are major control hardware failures. Recovery of safe flight is examined in this paper in the context of trajectory optimization. For a particular transport aircraft, and a failure scenario inspired by an historical air disaster, recovery scenarios are calculated with and without control surface failures, to bring the aircraft to safe flight from the adverse flight condition that it had assumed, apparently as a result of contact with a vortex from a larger aircraft's wake. An effort has been made to represent relevant airframe dynamics, acceleration limits, and actuator limits faithfully, since these contribute to the lack of agility and control power that plays an important role in defining what can be achieved with the vehicle when it is in extremis.

Sparks, Dean W.↗
NASA NTRS1979

Nonlinear decoupled control synthesis for maneuvering aircraft

A control law for decoupling roll rate, angle of attack and sideslip in rapid, nonlinear airplane maneuvers is derived. For simplicity, only moments caused by ailerons, rudder and elevator are considered, and control forces are neglected. Simulated responses of the closed loop system, including the control forces, show that the neglected forces have no significant effects and that large, simultaneous lateral and longitudinal maneuvers can be precisely performed. The violent divergences which occur in open-loop maneuvers of this type are eliminated by the decoupling. The control law showed little sensitivity to 20 percent perturbation in three important stability derivatives.

Singh, S. N.↗
NASA NTRS1974

The influence of wind shear on aerodynamic coefficients

The purpose of this study is to investigate the influence of wind shear on the lift, drag, roll and yaw moments of a wing in a horizontal wind gradient at various elevations and roll angles. The models of wind shear considered are those proposed by Leurs (1973) for atmospheric flow over horizontally homogeneous and uniform terrain. A general series solution for the distribution of lift along the wing span following the method of Houghton and Brock (1970) is employed to compute the loads and moments on the airfoil. Results of the computations indicate that wind shear can have a significant effect on the rolling and yawing moments of the wing of an aircraft flying with one wing low in the atmospheric boundary layer. These moments are directly influenced by the magnitude of the wind gradient in the neutral atmosphere reflected by the magnitude of the friction velocity. The effect of wind shear on lift and drag is negligible. Stability of the atmosphere tends to increase the magnitude of the rolling moment while decreasing its variation with elevation.

Frost, W.↗
NASA NTRS1948

Longitudinal Trim and Tumble Characteristics of a 0.057-Scale Model of the Chance Vought XF7U-1 Airplane, TED NO. NACA DE311

Based on results of longitudinal trim and tumble tests of a 0.057-scale model of the Chance Vought XF7U-1 airplane, the following conclusions regarding the trim and tumble characteristics of the airplane have been drawn: 1. The airplane will not trim at any unusual or uncontrolled angles of attack. 2. The airplane will not tumble with the center of gravity located forward of 24 percent of the mean aerodynamic chord. When the center of gravity is located at 24 percent of the mean aerodynamic chord and slats are extended and elevators are deflected full up, the airplane may tumble if given an external positive pitching moment. 3. The tumbling motion obtained will be readily terminated by deflecting the elevators full down so as to oppose the rotation. 4. The accelerations encountered during an established tumble may be dangerous to the pilot and, therefore, action should be taken to terminate a tumble immediately upon its inception. 5. Simultaneous opening of two wing-tip parachutes having diameters of 4 feet or larger and having drag coefficients of approximately 0.7 will effectively terminate the tumble. 6. Model results indicate that the pilot will not be struck by the airplane if it becomes necessary to leave the airplane during a tumble. The pilot may require aid from an ejection-seat arrangement.

Bryant, Robert L.↗
NASA NTRS1990

Wind-tunnel and flight-test investigation of the exdrone remotely piloted vehicle configuration

At the request of the United States Marine Corps, an exploratory wind-tunnel and flight test investigation was conducted by the NASA Langley Research Center to improve the stability, controllability, and general flight characteristics of the Marine Corps Exdrone RPV (Remotely Piloted Vehicle) configuration. Static wind tunnel tests were conducted to identify and improve the stability and control characteristics of the vehicle. The wind-tunnel test resulted in several configuration modifications which included increased elevator area, increased vertical tail area and moment arm, increased rudder area and aileron area, the addition of vertical wing-tip fins, and the addition of leading-edge droops on the outboard wing panel to improve the stall departure resistance. Flight tests of the modified configuration were conducted at the NASA Plum Tree Test Site to provide a qualitative evaluation of the flight characteristics of the modified configuration.

Yip, Long P.↗