ALBERT

Description: A second wind tunnel test of the FAST-MAC circulation control model was recently completed in the National Transonic Facility at the NASA Langley Research Center. The model was equipped with four onboard flow control valves allowing independent control of the circulation control plenums, which were directed over a 15% chord simple-hinged flap. The model was configured for low-speed high-lift testing with flap deflections of 30 and 60 degrees, along with the transonic cruise configuration with zero degree flap deflection. Testing was again conducted over a wide range of Mach numbers up to 0.88, and Reynolds numbers up to 30 million based on the mean chord. The first wind tunnel test had poor transonic force and moment data repeatability at mild cryogenic conditions due to inadequate thermal conditioning of the balance. The second test demonstrated that an improvement to the balance heating system significantly improved the transonic data repeatability, but also indicated further improvements are still needed. The low-speed highlift performance of the model was improved by testing various blowing slot heights, and the circulation control was again demonstrated to be effective in re-attaching the flow over the wing at off-design transonic conditions. A new tailored spanwise blowing technique was also demonstrated to be effective at transonic conditions with the benefit of reduced mass flow requirements.

Description: The Supersonics Project of the NASA Fundamental Aeronautics Program organized an internal sonic boom workshop to evaluate near- and mid-field sonic boom prediction capability at the Fundamental Aeronautics Annual Meeting in Atlanta, Georgia on October 8, 2008. Workshop participants computed sonic boom signatures for three non-lifting bodies and two lifting configurations. A cone-cylinder, parabolic, and quartic bodies of revolution comprised the non-lifting cases. The lifting configurations were a simple 69-degree delta wing body and a complete low-boom transport configuration designed during the High Speed Research Project in the 1990s with wing, body, tail, nacelle, and boundary layer diverter components. The AIRPLANE, Cart3D, FUN3D, and USM3D ow solvers were employed with the ANET signature propagation tool, output-based adaptation, and a priori adaptation based on freestream Mach number and angle of attack. Results were presented orally at the workshop. This article documents the workshop, results, and provides context on previously available and recently developed methods.

Description: A low-speed experimental investigation has been conducted on a 5.8-percent scale Hybrid Wing Body configuration in the NASA Langley 14- by 22-Foot Subsonic Tunnel. This Hybrid Wing Body (HWB) configuration was designed with specific intention to support the NASA Environmentally Responsible Aviation (ERA) Project goals of reduced noise, emissions, and fuel burn. This HWB configuration incorporates twin, podded nacelles mounted on the vehicle upper surface between twin vertical tails. Low-speed aerodynamic characteristics were assessed through the acquisition of force and moment, surface pressure, and flow visualization data. Longitudinal and lateral-directional characteristics were investigated on this multi-component model. The effects of a drooped leading edge, longitudinal flow-through nacelle location, vertical tail shape and position, elevon deflection, and rudder deflection have been studied. The basic configuration aerodynamics, as well as the effects of these configuration variations, are presented in this paper.

Description: In the past few years considerable improvement was made in NASA's in house boom prediction capability. As part of this improved capability, the USM3D Navier-Stokes flow solver, when combined with a suitable unstructured grid, went from accurately predicting boom signatures at 1 body length to 10 body lengths. Since that time, the research emphasis has shifted from analysis to the design of supersonic configurations with boom signature mitigation In order to design an aircraft, the techniques for accurately predicting boom and drag need to be determined. This paper compares CFD results with the wind tunnel experimental results conducted on a Gulfstream reduced boom and drag configuration. Two different wind-tunnel models were designed and tested for drag and boom data. The goal of this study was to assess USM3D capability for predicting both boom and drag characteristics. Overall, USM3D coupled with a grid that was sheared and stretched was able to reasonably predict boom signature. The computational drag polar matched the experimental results for a lift coefficient above 0.1 despite some mismatch in the predicted lift-curve slope.

Description: A novel multielement trailing-edge flap system for light general aviation airplanes was conceived for enhanced safety during normal and emergency landings. The system is designed to significantly reduce stall speed, and thus approach speed, with the goal of reducing maneuveringflight accidents and enhancing pilot survivability in the event of an accident. The research objectives were to assess the aerodynamic performance characteristics of the system and to evaluate the extent to which it provided both increased lift and increased drag required for the low-speed landing goal. The flap system was applied to a model of a light general aviation, high-wing trainer and tested in the Langley 12- Foot Low-Speed Wind Tunnel. Data were obtained for several device deflection angles, and component combinations at a dynamic pressure of 4 pounds per square foot. The force and moment data supports the achievement of the desired increase in lift with substantially increased drag, all at relatively shallow angles of attack. The levels of lift and drag can be varied through device deflection angles and inboard/outboard differential deflections. As such, it appears that this flap system may provide an enabling technology to allow steep, controllable glide slopes for safe rapid descent to landing with reduced stall speed. However, a simple flat-plate lower surface spoiler (LSS) provided either similar or superior lift with little impact on pitch or drag as compared to the proposed system. Higher-fidelity studies are suggested prior to use of the proposed system.

Description: Flight at speeds greater than the speed of sound is not permitted over land, primarily because of the noise and structural damage caused by sonic boom pressure waves of supersonic aircraft. Mitigation of sonic boom is a key focus area of the High Speed Project under NASA's Fundamental Aeronautics Program. The project is focusing on technologies to enable future civilian aircraft to fly efficiently with reduced sonic boom, engine and aircraft noise, and emissions. A major objective of the project is to improve both computational and experimental capabilities for design of low-boom, high-efficiency aircraft. NASA and industry partners are developing improved wind tunnel testing techniques and new pressure instrumentation to measure the weak sonic boom pressure signatures of modern vehicle concepts. In parallel, computational methods are being developed to provide rapid design and analysis of supersonic aircraft with improved meshing techniques that provide efficient, robust, and accurate on- and off-body pressures at several body lengths from vehicles with very low sonic boom overpressures. The maturity of these critical parallel efforts is necessary before low-boom flight can be demonstrated and commercial supersonic flight can be realized.

Description: A wind tunnel test and a computational study were conducted to investigate the complex interactions between a supersonic nozzle plume and shock waves of differing strengths generated from various aft surfaces typical of supersonic aircraft. These analytically-defined aft surfaces were representative of horizontal tails of various sizes, and an aft deck. CFD simulations of many proposed model configurations allowed for assessments of the detailed flow interactions of components in close proximity to the nozzle, as well as assessments of the nozzle jet flow itself. The evaluation of the computational results for many candidate configurations guided the design of model components. The interactions of the waveforms from these surfaces with the jet exhaust plume can have significant adverse effects on the loudness of the sonic boom if the surfaces are not carefully integrated into an aircraft design. The greatest discrepancy in estimating sonic boom loudness for low-boom flight vehicles is currently in predicting the signatures from the aft part of an aircraft, including the interactions with the plume flow. The objectives of this test were to gain a better understanding of these interactions, and to provide a detailed experimental database from multiple sources for use as validation cases for CFD tool development. The subject test was run in the NASA Ames 9- by 7-Ft Supersonic Wind Tunnel in February 2016 at Mach numbers of 1.6 and 2.0, and was funded by the NASA Commercial Supersonics Technology (CST) Project. The nozzle flow was provided by high-pressure air (HPA) pumped through the model, and pressure signature data were acquired with the NASA 14-inch sonic boom pressure rail. The rail measured the locations of the shocks and expansions at various distances and off-track angles from the model. This enabled the impact of the nozzle plume/shock interactions on the near- and mid-field sonic boom pressure waveforms to be quantified. Schlieren images of the flow field around and behind the model were obtained with an RBOS (Retroreflective Background-Oriented Schlieren) technique to determine the origins of the shock and expansion waves, to identify the shape and boundaries of the plume, and to determine the changes in incoming and exiting waveforms within the plume. A total pressure rake was positioned closely behind the model nozzle in order to measure the total pressure profiles of the flow above, within, and below the nozzle exhaust. Model angles and positions in the tunnel were measured by photogrammetry using two cameras since the lack of a model force balance prevented the measurement of model deflections under load.Navier-Stokes computations using two different CFD codes were compared to the experimental sonic boom pressure signature data, and the rake total pressure data in the plume. A computational schlieren technique was used to compare the computed flow field with the RBOS images. The computational results were also used to complement the test data with flow field quantities that could not be measured, such as Mach number and pressure distributions to distinguish shock waves and expansion waves.

Description: Near-field pressure signatures were measured and computational predictions made for several sonic boom models representing Boeings Quiet Experimental Validation Concept (QEVC) supersonic transport, as well as for three axisymmetric calibration models. The concept was designed under a NASA Research Announcement (NRA) contract to address environmental and performance goals, specifically for low sonic boom loudness levels and high cruise efficiency, for an aircraft anticipated to enter service in the 2020-timeframe. Wind tunnel tests were conducted on the aircraft and calibration models during Phases I and II of the NRA contract from 2011 to 2013 in the NASA Ames 9- by 7-Foot and NASA Glenn 8- by 6-Foot Supersonic Wind Tunnels. Sonic boom pressure signatures were acquired primarily at Mach 1.6 and 1.8, and force and moment data were acquired from Mach 0.8 to 1.8. The sonic boom test data were obtained using a 2-in. flat-top pressure rail and a 14-in. tapered reflection factor 1 (RF1) pressure rail. Both rails capture an entire pressure signature in one data point, and successive signatures at varying positions along or above the rail were used to improve data quality through spatial averaging. The sonic boom data obtained by the rails were validated with high-fidelity numerical simulations of offbody pressures. The test results showed good agreement between the computational and experimental data when a variety of testing techniques including spatial averaging of a series of pressure signatures were employed. The two wind tunnels generally produced comparable data.

Description: A computational and design study on an airfoil and high-lift flap for the X-57 Maxwell Distributed Electric Propulsion (DEP) testbed aircraft was conducted. The aircraft wing sizing study resulted in a wing area of 66.67 sq ft and aspect ratio of 15 with a design requirement of V(stall) = 58 KEAS, at a gross weight of 3,000 lb. To meet this goal an aircraft C(L,max) of 4.0 was required. The design cruise condition is 150 KTAS at 8,000 ft. This resulted in airfoil requirements of c(l) is approximately 0.90 for the cruise condition at Re = 2.35 x 10 (exp 6). A flapped airfoil with a c(l,max) of approximately 2.5 or greater, at Re = 1.0 x 10 (exp 6), was needed to have enough lift to meet the stall requirement with the DEP system. MSES computational analyses were conducted on the GAW-1, GAW-2, and the NACA 5415 airfoil sections, however they had limitations in either high drag or low c(l,max) on the cruise airfoil, which was the impetus for a new design. A design was conducted to develop a low drag airfoil for the X-57 cruise conditions with high c(l,max). The final design was the GNEW5BP93B airfoil with a minimum drag coefficient of c(d) = 0.0053 at c(l) = 0.90 and achieved laminar flow back to 69% chord on the upper surface and 62% chord on the lower surface. With fully turbulent flow, the drag increases to c(d) = 0.0120. The predicted maximum lift with turbulent flow is a c(l,max) of 1.95 at alpha = 19 deg. The airfoil is characterized by relatively flat pressure gradient regions on both surfaces at alpha = 0 deg, and aft camber to get extra lift out of the lower surface concave region. A 25% chord slotted flap was designed and analyzed with MSES for a 30 deg flap deflection. Additional 30 deg and 40 deg flap deflection analyses for two flap positions were conducted with USM3D using several turbulence models, for two angles of attack, to assess near c(l,max) with varied flap position. The maximum c(l) varied between 2.41 and 3.35. An infinite-span powered high-lift study was conducted on a GAW-1 constant chord 40 deg flapped airfoil section with FUN3D to quantify the airfoil lift increment that can be expected from a DEP system. The 16.7 hp/propeller blown wing increases the maximum C(L) from 3.45 to C(L) = 6.43, which is an effective q ratio of 1.86. This indicates that if the unblown high-lift flapped airfoil of the X-57 airplane achieves a c(l,max) of 2.78, then the high-lift augmentation blowing could yield a sectional lift coefficient of approximately 4.95 at c(l,max). Finally, a computational study was conducted with FUN3D on an infinite-span constant chord GAW-1 cruise airfoil to determine the impact of high-lift propeller diameter to wing chord ratio on the lift increment of the DEP system. A constant diameter propeller and nacelle size were used in the study. Three computational grids were made with airfoil chords of 0.5*chord, 1.0*chord, and 2.0*chord. Results of the propeller diameter to wing chord ratio study indicated that the blown to unblown C(L) ratio increased as the chord was decreased. However, because of the increase in relative size of the high-lift nacelle to the wing, which impacted wing lift performance, the study indicated that a propeller diameter to wing chord ratio of 1.0 gives the overall best maximum lift on the wing with the DEP system.

Description: A wind tunnel test was performed in the NASA Ames 9x7 Supersonic Wind Tunnel focusing on the shock waves traveling through and interacting with an exhaust nozzle plume. This experimental study was conducted to develop and validate the CFD capability required to accurately include nozzle flow with impinging shock effects on near field and groundpropagated sonic boom signatures. The model was made to be generic, and included a simple nozzle shape, two different aft decks, and a few generic horizontal tails. High pressure air was pumped through a nozzle at various nozzle pressure ratios (NPR) to represent the engine plume in flight. The three different aft body representations each created a different shock wave signature that passed through the plume. An aft deck configuration, where part of the aircraft shields the nozzle plume, was also tested. Retroreflective Background-Oriented Schlieren (RBOS) was used to obtain schlieren images of the flow field around the model and behind the model. This study compares wind tunnel data and numerical simulations conducted by the NASA Tetrahedral Unstructured Software System CFD code, USM3D.