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NASA NTRS1994

Controlling forebody asymmetries in flight: Experience with boundary layer transition strips

The NASA Dryden Flight Research Center has an ongoing program to investigate aircraft flight characteristics at high angles of attack. As part of this investigation, longitudinal boundary layer transition strips were installed on the F-18 HARV forebody, a preproduction F/A-18 radome with a nose-slice tendency, and the X-31 aircraft forebody and noseboom to reduce asymmetric yawing moments at high angles of attack. The transition strips were effective on the F-18 HARV at angles of attack above 60 deg. On the preproduction F/A-18 radome at an angle of attack near 50 deg the strips were not effective. When the transition strips were installed on the X-31 noseboom, a favorable effect was observed on the yawing moment dynamics but the magnitude of the yawing moment was not decreased.

David F Fisher↗

NASA NTRS1941

Comparison of Vee-Type and Conventional Tail Surfaces in Combination with Fuselage and Wing in the Variable-Density Tunnel

The pitching and the yawing moments of a vee-type and a conventional type of tail surface were measured. The tests were made in the presence of a fuselage and a wing-fuselage combination in such a way as to determine the moments contributed by the tail surfaces. The results showed that the vee-type tail tested, with a dihedral angle of 35.3 deg, was about 71 percent as effective in pitch as the conventional tail and had a yawing-moment to pitching-moment ratio of 0.3. The conventional tail, the panels of which were all congruent to those of the vee-type tail, had a yawing-moment to pitching-moment ratio of 0.48. These ratios are in fair agreement with values calculated by methods shown in this and previous reports. The values of the measured moments were reduced from 15 to 25 percent of the calculated value by fuselage interference.

Greenberg, Harry↗

NASA NTRS1954

Experimental investigation of the forces and moments due to sideslip of a series of triangular vertical and horizontal tail combinations at Mach numbers

An experimental investigation was made at Mach numbers of 1.62, 1.93, and 2.41 of a series of tail combinations consisting of triangular vertical tail attached symmetrically to a triangular horizontal tail to determine the lateral force, yawing moment, and rolling moment due to sideslip. The apex angles of both the vertical- and horizontal-tail surfaces were varied systematically in order to obtain results for an appreciable range of operating conditions. The results of the investigation indicated that, for tails having subsonic leading edges and supersonic trailing edges, the lateral-force derivative and the yawing-moment derivative were predicted satisfactorily by the method presented in NACA TN 3071 except when the leading edges approach a sonic condition. The theoretical rolling-moment derivative was in fair agreement with the experimental derivative. For the limited tests in which both the leading and trailing edges were supersonic, the prediction of the lateral-force derivative and the yawing-moment derivative obtained from NACA TN 2412 was in good agreement with the experimental derivatives, whereas the prediction of the rolling-moment derivative was fair.

Coletti, Donald E↗

NASA NTRS1993

Forebody vortex control on an F/A-18 using small, rotatable 'tip-strakes'

A mechanical scheme for manipulating the forebody vortices of an F/A-18, therefore creating controlled yawing moments at moderate and high angles of attack, was investigated in a wind tunnel. The technique consists of rotating miniature strakes (single or dual) about the radome centerline very close to the tip of the model. Forces, moments, and pressures were measured for angles of attack up to 60 deg and sideslip angles up to -10 deg. Results indicate that single and dual strakes can produce changes in side force and yawing moment, with magnitudes comparable to, or in some cases higher than, the directional changes produced by a 30-deg rudder deflection at 0-deg angle of attack. According to its circumferential position, the strake alters the separation location, inducing different degrees of asymmetry in the forebody vortex flow field that are translated into changes in side force and yawing moment. In comparison, the dual strakes appear to provide more gradual and better behaved changes than the single strake. Excellent correlation was found between this test and water tunnel tests performed on a similar configuration.

Suarez, Carlos J.↗

NASA NTRS1993

Subsonic aerodynamic characteristics of a proposed advanced manned launch system orbiter configuration

The Advanced Manned Launch System is a proposed near-term technology, two-stage, fully reusable launch system that consists of an unmanned glide-back booster and a manned orbiter. An orbiter model that featured a large fuselage and an aft delta wing with tip fins was tested in the Langley 7- by 10-Foot High-Speed Tunnel. A crew cabin, large payload fairing, and crew access tunnel were mounted on the upper body. The results of the investigation indicated that the configuration was longitudinally stable to an angle of attack of about 6 deg about a center-of-gravity position of 0.7 body length. The model had an untrimmed lift-drag ratio of 6.6, but could not be trimmed at positive lift. The orbiter model was also directionally unstable. The payload fairing was responsible for about half the instability. The tip-fin controllers, which are designed as active controls to produce artificial directional stability, were effective in producing yawing moment, but sizable adverse rolling moment occurred at angles of attack above 6 deg. Differential deflection of the elevon surfaces was effective in producing rolling moment with only small values of adverse yawing moment.

Ware, George M.↗

NASA NTRS1989

Low-speed, high-lift aerodynamic characteristics of slender, hypersonic accelerator-type configurations

Two investigations were conducted in the Langley 14 by 22 Foot Subsonic Tunnel to determine the low-speed aerodynamic characteristics of a generic hypersonic accelerator-type configuration. The model was a delta wing configuration incorporating a conical forebody, a simulated wrap-around engine package, and a truncated conical aftbody. Six-component force and moment data were obtained over a range of attack from -4 to 30 degrees and for a sideslip range of + or - 20 degrees. In addition to tests of the basic configuration, component build-up tests were conducted; and the effects of power, forebody nose geometry, canard surfaces, fuselage strakes, and engines on the lower surface alone were also determined. Control power available from deflections of wing flaps and aftbody flaps was also investigated and found to be significantly increased during power-on conditions. Large yawing moments resulted from asymmetric flow fields exhibited by the forebody as revealed by both surface pressure data and flow visualization. Increasing nose bluntness reduced the yawing-moment asymmetry, and the addition of a canard eliminated the yawing-moment asymmetry.

Gatlin, Gregory M.↗

NASA NTRS1949

Investigation at low speeds of the effect of aspect ratio and sweep on rolling stability derivatives of untapered wings

A low scale wind tunnel investigation was conducted in rolling flow to determine the effects of aspect ratio and sweep (when varied independently) on the rolling stability derivatives for a series of untapered wings. Test results indicate that when the aspect ratio was held constant, an increase in the sweepback angle caused a significant reduction in the damping in roll at low lift coefficients for only the higher aspect ratios that were tested. This result was in agreement with available swept wing theory which indicated no effect of sweep for aspect ratios near zero. The result of the linear theory that the damping in roll is independent of lift coefficient and that the yawing moment and lateral force due to rolling are directly proportional to the lift coefficient was found to be valid for only a very limited lift coefficient range when the wings were highly swept. For such wings, the damping was found to increase in magnitude and the yawing moment due to rolling, to change from negative to positive at moderate lift coefficients. The effect of wing tip suction, not acounted for by present theory, was found to be very important with regard to the yawing moment due to rolling, particularly for low aspect ratio swept wings. An empirical means of correcting present theory for the effect of tip suction is suggested.

Goodman, Alex↗

NASA NTRS1993

Aerodynamic control of NASP-type vehicles through Vortex manipulation. Volume 1: Static water tunnel tests

Water tunnel tests were conducted on a NASP-type configuration to evaluate different pneumatic Forebody Vortex Control (FVC) methods. Flow visualization and yawing moment measurements were performed at angles of attack from 0 deg to 30 deg. The pneumatic techniques tested included jet and slot blowing. In general, blowing can be used efficiently to manipulate the forebody vortices at angles of attack greater than 20 deg. These vortices are naturally symmetric up to alpha = 25 deg and asymmetric between 25 deg and 30 deg angle of attack. Results indicate that tangential aft jet blowing is the most promising method for this configuration. Aft jet blowing produces a yawing moment towards the blowing side and the trends with blowing rate are well behaved. The size of the nozzle is not the dominant factor in the blowing process; the change in the blowing 'momentum,' i.e., the product of the mass flow rate and the velocity of the jet, appears to be the important parameter in the water tunnel (incompressible and unchoked flow at the nozzle exit). Forward jet blowing is very unpredictable and sensitive to mass flow rate changes. Slot blowing (with the exception of very low blowing rates) acts as a flow 'separator'; it promotes early separation on the blow side, producing a yawing moment toward the non-blowing side for the C(sub mu) range investigated.

Suarez, Carlos J.↗

NASA NTRS1937

Effect of tip shape and dihedral on lateral-stability characteristics

This report presents the results of wind tunnel tests to determine the effect of wing-tip shape and dihedral on some of the aerodynamic characteristics of Clark Y wings that affect the performance and lateral stability of airplanes. Force tests at several angles of yaw and rotation tests at zero yaw were made. From these tests the rates of change of rolling moment, yawing moment, and cross-wind force coefficients with angle of yaw and the rate of change of rolling moment coefficient with rolling were determined.

Shortal, Joseph A↗

NASA NTRS1993

Subsonic aerodynamic characteristics of the HL-20 lifting-body configuration

The HL-20 is proposed as a possible future manned spacecraft. The configuration consists of a low-aspect-ratio body with a flat undersurface. Three fins (a small centerline fin and two outboard (tip) fins set at a dihedral angle of 50 deg) are mounted on the aft body. The control system consists of elevon surfaces on the outboard fins, a set of four body flaps on the upper and lower aft body, and an all-movable center fin. Both the elevons and body flaps were capable of trimming the model to angles of attack from -2 deg to above 20 deg. The maximum trimmed lift-drag ratio was 3.6. Replacing the flat-plate tip fins with airfoil tip fins increased the maximum trimmed lift-drag ratio to 4.2. The elevons were effective as a roll control, but they produced about as much yawing moment as rolling moment because of the tip-fin dihedral angle. The body flaps produced less rolling moment than the elevons and only small values of yawing moment. A limited investigation of the effect of varying tip-fin dihedral angle indicated that a dihedral angle of 50 deg was a reasonable compromise for longitudinal and lateral stability, longitudinal trim, and performance at subsonic speeds.

Ware, George M.↗

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 NTRS1985

Filament wound data base development, revision 1, appendix A

Data are presented in tabular form for the High Performance Nozzle Increments, Filament Wound Case (FWC) Systems Tunnel Increments, Steel Case Systems Tunnel Increments, FWC Stiffener Rings Increments, Steel Case Stiffener Rings Increments, FWC External Tank (ET) Attach Ring Increments, Steel Case ET Attach Ring Increments, and Data Tape 8. The High Performance Nozzle are also presented in graphical form. The tabular data consist of six-component force and moment coefficients as they vary with angle of attack at a specific Mach number and roll angle. The six coefficients are normal force, pitching moment, side force, yawing moment, axial force, and rolling moment. The graphical data for the High Performance Nozzle Increments consist of a plot of a coefficient increment as a function of angle of attack at a specific Mach number and at a roll angle of 0 deg.

Sharp, R. Scott↗

NASA NTRS1985

Filament wound data base development, revision 1, appendix B

The data tape 8 data are presented in tabular form. Each page consists of six-component force and moment coefficients as they vary with angle of attack at a specific Mach number and roll angle. The angle of attack range is from 0 to 180 deg in 5 deg increments. The six coefficients are normal force, pitching moment, side force, yawing moment, axial force, and rolling moment. Each successive page is incremented for roll angle in 45 deg increments for a roll angle range from 0 to 360 deg. The cycle is then repeated for the next Mach number. The Mach numbers in the data base are 0.4, 0.5, 0.55, 0.6, 0.7, 0.8, 0.9, 0.95, 1.05, 1.1, 1.2, 1.4, 1.96, 2.5, 3.0, and 3.48.

Sharp, R. Scott↗

NASA NTRS1946

Jet-boundary and Plan-form Corrections for Partial-Span Models with Reflection-Plane, End-Plate, or No End-Plate in a Closed Circular Wind Tunnel

A method is presented for determining the jet-boundary and plan-form corrections necessary for application to test data for a partial-span model with a reflection plane, an end plate, or no end plate in a closed circular wind tunnel. Examples are worked out for a partial-span model with each of the three end conditions in the Langley 19-foot pressure tunnel and the corrections are applied to measured values of lift, drag, pitching-moment, rolling-moment, and yawing-moment coefficients.

Sivells, James C↗

NASA NTRS1947

Lateral Stability Characteristics of a 1/8.33-Scale Powered Model of the Republic XF-12 Airplane

The XF-12 airplane is a high-performance photo-reconnaissance aircraft designed for the Army Air Forces by the Republic Aviation Corporation. An investigation of a 1/8.33 - scale powered model was made in the Langley l9-foot pressure tunnel to obtain information relative to the aerodynamic design of the airplane. The model was tested with and without the original vertical tail. and with two revised tails. For the revised tail no. 1, the span of the original vertical .tail was increased about 15 percent and the portion of the vertical tail between the stabilizer and fuselage behind the rudder hinge line was allowed to deflect simultaneously with the main rudder. Revision no. 2 incorporated the increased span, but the lower rudder was locked in the neutral position. For all the tail arrangements investigated it was indicated that the airplane will possess positive effective dihedral and will be directionally stable regardless of flap or power condition. The rudder effectiveness is greater for the revised tails than for the original tail, but this is offset by the increase in directional stability caused by the revised tail. All the rudder arrangements appear inadequate in trimming out the resultant yawing moments at zero yaw in a take - off condition with the left-hand outboard propeller windmilling and the remaining engines developing take-off power.

Pepper, Edward↗

NASA NTRS1995

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.↗

NASA NTRS1995

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

Pedreiro, N.↗

NASA NTRS1939

Pressure-distribution measurements on a rectangular wing with a partial-span split flap in curved flight

Pressure-distribution tests were made on the 32-foot whirling arm of the Daniel Guggenheim Airship Institute of a rectangular wing of NACA 23012 section to determine the rolling and the yawing moment due to angular velocity in yaw. The model was tested at 0 and 5 degree pitch; 0, +/- 5, and +/- 10 degrees yaw; and with no flap and with split flaps 25, 50 and 75 percent of the wing span and deflected 60 degrees. The results are given in the form of span load distributions and as calculated moment coefficients. The experimental values of rolling- and yawing-moment coefficients were in fairly close agreement with theory.

Rokus, Frank G↗