National Full-Scale Aerodynamics Complex Research Articles

Active Flow Control on Vertical Tail Models

Active flow control (AFC) subscale experiments were conducted at the Lucas Wind Tunnel of the California Institute of Technology. Tests were performed on a generic vertical tail model at low speeds. Fluidic oscillators were used at the trailing edge of the main element (vertical stabilizer) to redirect the flow over the rudder and delay or prevent flow separation. Side force increases in excess of 50% were achieved with a 2% momentum coefficient () input. The results indicated that a collective of about 1% could increase the side force by 30–50%. This result is achieved by reducing the spanwise flow on the swept back wings that contributes to early flow separation near their tips. These experiments provided the technical backdrop to test the full-scale Boeing 757 vertical tail model equipped with a fluidic oscillator system at the National Full-scale Aerodynamics Complex 40-by 80-foot Wind Tunnel, NASA Ames Research Center. The is shown to be an important parameter for scaling a fluidic oscillator AFC system from subscale to full-scale wind tunnel tests. The results of these tests provided the required rationale to use a fluidic oscillator AFC configuration for a follow-on flight test on the Boeing 757 ecoDemonstrator.

Full-Scale Testing of Active Flow Control Applied to a Vertical Tail

Full-scale wind tunnel tests were carried out on a Boeing 757 vertical tail equipped with 37 sweeping jet actuators on the starboard side along the rudder hinge line. The tests were performed at the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel () at NASA Ames Research Center. The model was tested at a nominal airspeed of 100 knots across rudder deflections and sideslip angles that covered the airplane’s emergency single engine climb. Active flow control (AFC) was optimized at the maximum rudder deflection of 30° and sideslip angles of 0° and . A threshold success criterion of 20% increase in the maximum side force was exceeded at largest rudder deflection using 12 sweeping jet actuators in the absence of sideslip but barely reached it with 18 actuators at the maximum sideslip at a momentum coefficient input of approximately 0.5%; consequently, 31 actuators were selected for most tests. AFC caused significant increases in suction pressure and associated side force on the actuated side. The successful demonstration of this application cleared the way for a subsequent flight demonstration on the Boeing 757 ecoDemonstrator in 2015.

Wind Tunnel Test of the SMART Active Flap Rotor

Boeing and a team from Air Force, NASA, Army, DARPA, MIT, UCLA, and U. of Maryland have successfully completed a wind-tunnel test of the smart material actuated rotor technology (SMART) rotor in the 40- by 80-foot wind-tunnel of the National Full-Scale Aerodynamic Complex at NASA Ames Research Center. The Boeing SMART rotor is a full-scale, five-bladed bearingless MD 900 helicopter rotor modified with a piezoelectric-actuated trailing edge flap on each blade. The eleven-week test program evaluated the forward flight characteristics of the active-flap rotor at speeds up to 155 knots, gathered data to validate state-of-the-art codes for rotor aero-acoustic analysis, and quantified the effects of open and closed loop active flap control on rotor loads, noise, and performance. The test demonstrated on-blade smart material control of flaps on a full-scale rotor for the first time in a wind tunnel. The effectiveness of the active flap control on noise and vibration was conclusively demonstrated. Results showed significant reductions up to 6dB in blade-vortex-interaction and in-plane noise, as well as reductions in vibratory hub loads up to 80%. Trailing-edge flap deflections were controlled within 0.1 degrees of the commanded value. The impact of the active flap on control power, rotor smoothing, and performance was also demonstrated. Finally, the reliability of the flap actuation system was successfully proven in more than 60 hours of wind-tunnel testing.

Thermal Protection System Response Uncertainty of a Hypersonic Inflatable Aerodynamic Decelerator

The objective of this paper is to investigate the uncertainty in the bondline temperature response of a flexible thermal protection system subject to uncertain parameters in the hypersonic flowfield and thermal response modeling of a hypersonic inflatable aerodynamic decelerator configuration for ballistic Mars entry. An inflatable decelerator with a 10 m major diameter is selected for this study based on the forebody dimensions scaled from the 6 m test article that was tested in the NASA Ames Research Center’s National Full-Scale Aerodynamics Complex facility. A global nonlinear sensitivity analysis study for the bondline temperature uncertainty shows that the dimension of uncertain parameters can reduce from 22 to 8. An uncertainty analysis of the bondline temperature in the reduced dimensions indicates that the bondline temperature varies by as much as 125% above the nominal prediction and exceeds the temperature limit of 400°C. The largest uncertainty occurred at 70 s in the trajectory before separation of the inflatable decelerator for transition to a secondary descent technology. The main contributors to the bondline temperature uncertainty are the insulator and outer fabric conductivities, as well as the freestream density. The thickness and initial density of the insulator layer, closest to the gas barrier layer, are also shown to be significant contributors to the bondline temperature uncertainty, especially earlier in the trajectory.

Airloads Correlation of the UH-60A Rotor inside the 40- by 80-Foot Wind Tunnel

The presented research validates the capability of a loosely coupled computational fluid dynamics (CFD) and comprehensive rotorcraft analysis (CRA) code to calculate the flowfield around a rotor and test stand mounted inside a wind tunnel. The CFD/CRA predictions for the Full-Scale UH-60A Airloads Rotor inside the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel at NASA Ames Research Center are compared with the latest measured airloads and performance data. The studied conditions include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed up to and including stall. For the speed sweep, wind tunnel modeling becomes important at advance ratios greater than 0.37 and test stand modeling becomes increasingly important as the advance ratio increases. For the thrust sweep, both the wind tunnel and test stand modeling become important as the rotor approaches stall. Despite the beneficial effects of modeling the wind tunnel and test stand, the new models do not completely resolve the current airload discrepancies between prediction and experiment.

FUN3D Airloads Predictions for the Full-Scale UH-60A Airloads Rotor in a Wind Tunnel

An unsteady Reynolds-Averaged Navier-Stokes solver for unstructured grids, FUN3D, is used to compute the rotor performance and airloads of the UH-60A Airloads Rotor in the National Full-Scale Aerodynamic Complex (NFAC) 40- by 80-foot Wind Tunnel. The flow solver is loosely coupled to a rotorcraft comprehensive code, CAMRAD-II, to account for trim and aeroelastic deflections. Computations are made for the 1-g level flight speed-sweep test conditions with the airloads rotor installed on the NFAC Large Rotor Test Apparatus (LRTA) and in the 40- by 80-ft wind tunnel to determine the influence of the test stand and wind-tunnel walls on the rotor performance and airloads. Detailed comparisons are made between the results of the CFD/CSD simulations and the wind tunnel measurements. The computed trends in solidity-weighted propulsive force and power coefficient match the experimental trends over the range of advance ratios and are comparable to previously published results. Rotor performance and sectional airloads show little sensitivity to the modeling of the wind-tunnel walls, which indicates that the rotor shaft-angle correction adequately compensates for the wall influence up to an advance ratio of 0.37. Sensitivity of the rotor performance and sectional airloads to the modeling of the rotor with the LRTA body/hub increases with advance ratio. The inclusion of the LRTA in the simulation slightly improves the comparison of rotor propulsive force between the computation and wind tunnel data but does not resolve the difference in the rotor power predictions at mu = 0.37. Despite a more precise knowledge of the rotor trim loads and flight condition, the level of comparison between the computed and measured sectional airloads/pressures at an advance ratio of 0.37 is comparable to the results previously published for the high-speed flight test condition.

Development and Operation of an Automatic Rotor Trim Control System for the UH-60 Individual Blade Control Wind Tunnel Test

A full-scale wind tunnel test to evaluate the effects of Individual Blade Control (IBC) on the performance, vibration, noise and loads of a UH-60A rotor was recently completed in the National Full-Scale Aerodynamics Complex (NFAC) 40- by 80-Foot Wind Tunnel [1]. A key component of this wind tunnel test was an automatic rotor trim control system that allowed the rotor trim state to be set more precisely, quickly and repeatably than was possible with the rotor operator setting the trim condition manually. The trim control system was also able to maintain the desired trim condition through changes in IBC actuation both in open- and closed-loop IBC modes, and through long-period transients in wind tunnel flow. This ability of the trim control system to automatically set and maintain a steady rotor trim enabled the effects of different IBC inputs to be compared at common trim conditions and to perform these tests quickly without requiring the rotor operator to re-trim the rotor. The trim control system described in this paper was developed specifically for use during the IBC wind tunnel test

Long-range Schmidt-Cassegrain laser velocimeter for large wind-tunnel applications

A two-component dual-beam laser velocimeter was designed and fabricated for use in the NASA Ames 40 X 80 ft wind tunnel. The instrument design was predicated on the use of inexpensive, high-quality, large-aperture Schmidt-Cassegrain telescope optics to perform both the laser beam focusing and scattered light collection functions. The instrument was configured to measure the streamwise and cross-stream components of velocity and enclosed in a streamlined optical assembly mounted on a two-dimensional traverse on the floor of the test section. The laser velocimeter probe volume was translated in the vertical direction by means of zoom optics. Coaxial backscattered light was collected through the same optical system and analyzed for Doppler shift. Single-mode polarization-maintaining fiber optics transmitted light from the remotely located laser into the optical assembly. A multimode fiber conveyed the scattered light to remotely located photodetectors. Standard frequency domain processing permitted analysis of Doppler signals with very poor signal-to-noise ratio. The system was evaluated in an optics laboratory in preparation for wind-tunnel test program applications. Sample data from this evaluation were presented as representative of the capability of the instrument. Statement of the Problem I N recent decades the laser Doppler velocimeter1 (LV) has seen increasing application to the study of aerodynamic phenomena. The LV provides a noninvasive measurement capability that is critical for the study of many phenomena. The dual-beam or differentialDoppler configuration of the instrument2 has been shown to have many practical advantages and is used extensively in small-scale wind-tunnel testing. Although several new techniques offer increased potential for global measurement capability,3'4 the LV remains a very important instrument for any aeronautical research facility. Although significant progress has been made,5 LV has not been widely used in large-scale wind-tunnel testing. To maintain the noninvasive quality of the measurement, it is necessary to place the instrument outside the wind-tunnel test section, or at least far enough away to limit its effect on the flowfield of interest. In largescale wind tunnels, this means large separation between the probe volume (laser beam intersection volume) where Doppler shifted light is scattered and the receiving optics where that light is collected and focused on a photodetector. This large separation or longrange operation requirement demands high laser power density in the probe volume to achieve high scattered intensity of the Doppler shifted light and a large collection solid angle to concentrate it on the photodetector. For a closed jet wind tunnel, it is further necessary to use windows of sufficient quality that the laser beams launched into the test section as well as the scattered light collected from the probe volume are not significantly degraded. In particular, the laser beams must not be refracted by striae in the windows so as to significantly reduce beam overlap in the probe volume. The window aperture must be sufficient to access the entire flow region of interest. The National Full-Scale Aerodynamics Complex at NASA Ames Research Center is composed of two large wind tunnels and a static (hover) test facility. The 40 x 80 ft wind tunnel is the smaller of the two wind tunnels, with a speed capability up to 300 kn (154 m/s).

Propulsion-induced aerodynamic effects measured with a full-scale STOVL model

Propulsion-induced aerodynamic interference effects are presented for a full-scale, powered, STOVL fighter aircraft model. The ejector-lift/vectored thrust configuration, designated the E-7A, was tested at low speed in the 40by 80and 80by 120-ft wind tunnels of the National Full-Scale Aerodynamics Complex located at NASA Ames Research Center. Aerodynamic effects on vehicle lift, drag, and pitching moment are presented over a range of effective velocities for simultaneous operation of all lifting jets. The jet/airframe interactions for separate operation of the lifting ejector system and vector able ventral nozzle are also presented. Ejector and engine inlet momentum effects were fully simulated in these full-scale, powered tests. The jet thrust vector angle of the ventral nozzle was varied to simulate transition from hover to wingborne flight modes. When the lifting ejector system and ventral nozzle are operated simultaneously, the induced effects on lift decrease as the thrust vector angle of the ventral nozzle approaches the horizontal. A negative increment in drag is produced over a narrow portion of the transition speed range when the ejectors and ventral nozzle are operated together. Aerodynamic induced effects of the ejector system measured at full-scale compare well with the small-scale data. Changes in lift and pitching moment due to ventral nozzle operation are smaller at full scale.

Comparison of model- and full-scale wind-tunnel performance

Experimental results from the Integrated Systems Test of the upgraded National Full-Scale Aerodynamics Complex 40- by 80-ft wind tunnel located at NASA Ames Research Center, which took place in 1987, are compared with results obtained with a 1/50 scale model of the wind tunnel. Test-section flow characteristics, air-exchange performance, total pressure distributions, and wall-static pressure distributions obtained with the model- and full-scale facilities are compared. The 1/50-scale model data were used to predict qualitative performance trends of the full-scale facility during investigation of design changes. Despite large differences in scale and Reynolds number between the two tunnels, quantitative agreement between certain model and full-scale performance characteristics was quite good.