ALBERT

Description: This work is a simulation technology demonstrator, of sweep jets used to suppress boundary layer separation and increase maximum achievable load coefficients. A sweep jet is a discrete Coanda jet that oscillates in the plane parallel to an aerodynamic surface. It injects mass and momentum in the approximate stream wise direction. It also generate turbulent eddies at the oscillation frequency, which are typically large relative to boundary layer turbulence, and which augmenting mixing across the boundary layer to attack flow separation. Simulations of a fluidic oscillator, the sweep jet emerging from the oscillator, and the suppression of boundary layer separation by an array of sweep jets are performed. Simulation results are compared to data from a dedicated CFD validation experiment of a single oscillator and its sweep jet, and from a study of a full-scale Boeing 757 vertical tail augmented with an array of sweep jets.2, 20 A critical step in the work is the development of realistic time-dependent sweep-jet in flow boundary conditions, derived from the results of the single-oscillator simulations, which create the sweep jets in the full-tail simulations. Simulations were performed using the Over flow CFD solver, with high-order spatial discretization and a range of turbulence modeling. Good results were obtained for all flows simulated, when suitable turbulence modeling was used.

Description: A series of aeroelastic optimization problems are solved on a high aspect ratio wingbox of the Common Research Model, in an effort to minimize structural mass under coupled stress, buckling, and flutter constraints. Two technologies are of particular interest: tow steered composite laminate skins and curvilinear stiffeners. Both methods are found to afford feasible reductions in mass over their non-curvilinear structural counterparts, through both distinct and shared mechanisms for passively controlling aeroelastic performance. Some degree of diminishing returns are seen when curvilinear stiffeners and curvilinear fiber tow paths are used simultaneously.

Description: An overview of aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign test vehicle is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a helium balloon, then accelerating the TV to Mach 4 and 53 km altitude with a solid rocket motor. Test flights conducted in June of 2014 (SFDT-1) and 2015 (SFDT-2) each successfully delivered a 6 meter diameter decelerator (SIAD-R) to test conditions and several seconds of flight, and were successful in demonstrating the SFDT flight system concept and SIAD-R technology. Aerodynamic models and uncertainties developed for the SFDT campaign are presented, including the methods used to generate them and their implementation within an aerodynamic database (ADB) routine for flight simulations. Pre- and post-flight aerodynamic models are compared against reconstructed flight data and model changes based upon knowledge gained from the flights are discussed. The pre-flight powered phase model is shown to have a significant contribution to off-nominal SFDT trajectory lofting, while coast and SIAD phase models behaved much as predicted.

Description: No abstract available

Description: The aerodynamic effects of compliant flaps installed onto a modified Gulfstream III airplane were investigated. Analyses were performed prior to flight to predict the aerodynamic effects of the flap installation. Flight tests were conducted to gather both structural and aerodynamic data. The airplane was instrumented to collect vehicle aerodynamic data and wing pressure data. A leading-edge stagnation detection system was also installed. The data from these flights were analyzed and compared with predictions. The predictive tools compared well with flight data for small flap deflections, but differences between predictions and flight estimates were greater at larger deflections. This paper describes the methods used to examine the aerodynamics data from the flight tests and provides a discussion of the flight-test results in the areas of vehicle aerodynamics, wing sectional pressure coefficient profiles, and air data.

Description: Computational fluid dynamics (CFD) analysis was conducted to study the low-speed stall aerodynamics of a Gulfstream G-III airplane (Gulfstream Aerospace Corporation, Savannah, Georgia) swept wing modified with an experimental seamless, compliant flap called the Adaptive Compliant Trailing Edge (ACTE) flap. The stall characteristics of the modified ACTE wing were analyzed and compared with the unmodified, clean wing at the flight speed of 120 knots and altitude of 2300 feet above mean sea level, in free air as well as in ground effect. A polyhedral finite-volume unstructured full Navier-Stokes CFD code, STAR-CCM (registered trademark) plus (CD-adapco [Computational Dynamics Limited, United Kingdom, and Analysis & Design Application Co., United States]), was used. Steady Reynolds-averaged Navier-Stokes CFD simulations were conducted for a clean wing and the ACTE wings at various ACTE deflection angles in free air (-2 degrees, 15 degrees, and 30 degrees) as well as in ground effect (15 degrees and 30 degrees). Solution sensitivities to grid densities were examined. In free air, the ACTE wings are predicted to stall at lower angles of attack than the clean wing. In ground effect, all wings are predicted to stall at lower angles of attack than the corresponding wings in free air. Even though the lift curves are higher in ground effect than in free air, the maximum lift coefficients for all wings are lower in ground effect. Finally, the lift increase due to ground effect for the ACTE wing is predicted to be less than the clean wing.

Description: The NASA Environmentally Responsible Aviation (ERA) Project sponsored a series of computational and experimental investigations of the propulsion and airframe integration issues associated with Hybrid-Wing-Body (HWB) or Blended-Wing-Body (BWB) configurations. NASA collaborated with Boeing Research and Technology (BR&T) to conduct this research on a new twin-engine Boeing BWB transport configuration. The experimental investigations involved a series of wind tunnel tests with a 5.75-percent scale model conducted in two low-speed wind tunnels. This testing focused on the basic aerodynamics of the configuration and selection of the leading edge Krueger slat position for takeoff and landing. This paper reviews the results and analysis of these low-speed wind tunnel tests.

Description: A concerted effort has been underway over the past several years to evolve computational capabilities for modeling aircraft loss-of-control under the NASA Aviation Safety Program. A principal goal has been to develop reliable computational tools for predicting and analyzing the non-linear stability & control characteristics of aircraft near stall boundaries affecting safe flight, and for utilizing those predictions for creating augmented flight simulation models that improve pilot training. Pursuing such an ambitious task with limited resources required the forging of close collaborative relationships with a diverse body of computational aerodynamicists and flight simulation experts to leverage their respective research efforts into the creation of NASA tools to meet this goal. Considerable progress has been made and work remains to be done. This paper summarizes the status of the NASA effort to establish computational capabilities for modeling aircraft loss-of-control and offers recommendations for future work.

Description: The Columbia Scientific Balloon Facility provides Telemetry and Command systems necessary for balloon operations and science support. There are various Line-Of-Sight (LOS) and Over-The-Horizon (OTH) systems and interfaces that provide communications to and from a science payload. This presentation will discuss the current data throughput options available and future capabilities that may be incorporated in the LDB Support Instrumentation Package (SIP) such as doubling the TDRSS data rate. We will also explore some new technologies that could potentially expand the data throughput of OTH communications.

Description: Blade tip vortices generated by a helicopter rotor blade are a major source of rotor noise and airframe vibration. This occurs when a vortex passes closely by, and interacts with, a rotor blade. The accurate prediction of Blade Vortex Interaction (BVI) continues to be a challenge for Computational Fluid Dynamics (CFD). Though considerable research has been devoted to BVI noise reduction and experimental techniques for measuring the blade tip vortices in a wind tunnel, there are only a handful of post-processing tools available for extracting vortex core lines from CFD simulation data. In order to calculate the vortex core radius, most of these tools require the user to manually select a vortex core to perform the calculation. Furthermore, none of them provide the capability to track the growth of a vortex core, which is a measure of how quickly the vortex diffuses over time. This paper introduces an automated approach for tracking the core growth of a blade tip vortex from CFD simulations of rotorcraft in hover. The proposed approach offers an effective method for the quantification and visualization of blade tip vortices in helicopter rotor wakes. Keywords: vortex core, feature extraction, CFD, numerical flow visualization

Description: The Ames Vertical Gun Range (AVGR) is a national facility for conducting laboratory- scale investigations of high-speed impact processes. It provides a set of light-gas, powder, and compressed gas guns capable of accelerating projectiles to speeds up to 7 km s(exp -1). The AVGR has a unique capability to vary the angle between the projectile-launch and gravity vectors between 0 and 90 deg. The target resides in a large chamber (diameter approximately 2.5 m) that can be held at vacuum or filled with an experiment-specific atmosphere. The chamber provides a number of viewing ports and feed-throughs for data, power, and fluids. Impacts are observed via high-speed digital cameras along with investigation-specific instrumentation, such as spectrometers. Use of the range is available via grant proposals through any Planetary Science Research Program element of the NASA Research Opportunities in Space and Earth Sciences (ROSES) calls. Exploratory experiments (one to two days) are additionally possible in order to develop a new proposal.

Description: The Transonic Dynamics Tunnel (TDT) at the National Aeronautics and Space Administration's (NASA) Langley Research Center began research operations in early 1960. Since that time, over 600 tests have been conducted, primarily in the discipline of aeroelasticity. This paper presents a bibliography of the publications that contain data from these tests along with other reports that describe the facility, its capabilities, testing techniques, and associated research equipment. The bibliography is divided by subject matter into a number of categories. An index by author's last name is provided.

Description: A wing/fuselage wind-tunnel model was tested in the Langley 14- by 22-foot Subsonic Wind Tunnel in preparation for a highly-instrumented Juncture Flow Experiment to be conducted in the same facility. This test, which was sponsored by the NASA Transformational Tool and Technologies Project, is part of a comprehensive set of experimental and computational research activities to develop revolutionary, physics-based aeronautics analysis and design capability. The objectives of this particular test were to examine the surface and off-body flow on a generic wing/body combination to: 1) choose a final wing for a future, highly instrumented model, 2) use the results to facilitate unsteady pressure sensor placement on the model, 3) determine the area to be surveyed with an embedded laser-doppler velocimetry (LDV) system, 4) investigate the primary juncture corner- flow separation region using particle image velocimetry (PIV) to see if the particle seeding is adequately entrained and to examine the structure in the separated region, and 5) to determine the similarity of observed flow features with those predicted by computational fluid dynamics (CFD). This report documents the results of the above experiment that specifically address the first three goals. Multiple wing configurations were tested at a chord Reynolds number of 2.4 million. Flow patterns on the surface of the wings and in the region of the wing/fuselage juncture were examined using oil- flow visualization and infrared thermography. A limited number of unsteady pressure sensors on the fuselage around the wing leading and trailing edges were used to identify any dynamic effects of the horseshoe vortex on the flow field. The area of separated flow in the wing/fuselage juncture near the wing trailing edge was observed for all wing configurations at various angles of attack. All of the test objectives were met. The staff of the 14- by 22-foot Subsonic Wind Tunnel provided outstanding support and delivered exceptional value to the experiment, which exceeded expectations. The results of this test will directly inform the planning for the first of a series of instrumented-model tests at the same Reynolds number. These tests will be performed on a slightly larger-scale model with the selected wing, and will include off-body measurements with LDV and PIV, steady and unsteady pressure measurements, and the flow-visualization techniques that are discussed in this report.

Description: As part of a computational study of acoustic radiation due to the passage of turbulent boundary layer eddies over the trailing edge of an airfoil, the Lattice-Boltzmann method is used to perform direct numerical simulations of compressible, low Mach number flow past an NACA 0012 airfoil at zero degrees angle of attack. The chord Reynolds number of approximately 0.657 million models one of the test conditions from a previous experiment by Brooks, Pope, and Marcolini at NASA Langley Research Center. A unique feature of these simulations involves direct modeling of the sand grain roughness on the leading edge, which was used in the abovementioned experiment to trip the boundary layer to fully turbulent flow. This report documents the findings of preliminary, proof-of-concept simulations based on a narrow spanwise domain and a limited time interval. The inclusion of fully-resolved leading edge roughness in this simulation leads to significantly earlier transition than that in the absence of any roughness. The simulation data is used in conjunction with both the Ffowcs Williams-Hawkings acoustic analogy and a semi-analytical model by Roger and Moreau to predict the farfield noise. The encouraging agreement between the computed noise spectrum and that measured in the experiment indicates the potential payoff from a full-fledged numerical investigation based on the current approach. Analysis of the computed data is used to identify the required improvements to the preliminary simulations described herein.

Description: The effect of nonlinear optimal streaks on disturbance growth in a Mach 6 axisymmetric flow over a 7deg half-angle cone is investigated in an e ort to expand the range of available techniques for transition control. Plane-marching parabolized stability equations are used to characterize the boundary layer instability in the presence of azimuthally periodic streaks. The streaks are observed to stabilize nominally planar Mack mode instabilities, although oblique Mack mode disturbances are destabilized. Experimentally measured transition onset in the absence of any streaks correlates with an amplification factor of N = 6 for the planar Mack modes. For high enough streak amplitudes, the transition threshold of N = 6 is not reached by the Mack mode instabilities within the length of the cone, but subharmonic first mode instabilities, which are destabilized by the presence of the streaks, reach N = 6 near the end of the cone. These results suggest a passive flow control strategy of using micro vortex generators to induce streaks that would delay transition in hypersonic boundary layers.

Description: This report documents the data collected during the large wind tunnel campaigns conducted as part of the SUNSET project (StUdies oN Scaling EffecTs due to ice) also known as the Ice-Accretion Aerodynamics Simulation study: a joint effort by NASA, the Office National d'Etudes et Recherches Arospatiales (ONERA), and the University of Illinois. These data form a benchmark database of full-scale ice accretions and corresponding ice-contaminated aerodynamic performance data for a two-dimensional (2D) NACA 23012 airfoil. The wider research effort also included an analysis of ice-contaminated aerodynamics that categorized ice accretions by aerodynamic effects and an investigation of subscale, low- Reynolds-number ice-contaminated aerodynamics for the NACA 23012 airfoil. The low-Reynolds-number investigation included an analysis of the geometric fidelity needed to reliably assess aerodynamic effects of airfoil icing using artificial ice shapes. Included herein are records of the ice accreted during campaigns in NASA Glenn Research Center's Icing Research Tunnel (IRT). Two different 2D NACA 23012 airfoil models were used during these campaigns; an 18-in. (45.7-cm) chord (subscale) model and a 72-in. (182.9-cm) chord (full-scale) model. The aircraft icing conditions used during these campaigns were selected from the Federal Aviation Administration's (FAA's) Code of Federal Regulations (CFR) Part 25 Appendix C icing envelopes. The records include the test conditions, photographs of the ice accreted, tracings of the ice, and ice depth measurements. Model coordinates and pressure tap locations are also presented. Also included herein are the data recorded during a wind tunnel campaign conducted in the F1 Subsonic Pressurized Wind Tunnel of ONERA. The F1 tunnel is a pressured, high- Reynolds-number facility that could accommodate the full-scale (72-in. (182.9-cm) chord) 2D NACA 23012 model. Molds were made of the ice accreted during selected test runs of the full-scale model in the IRT. From these molds, castings were made that closely replicated the features of the accreted ice. The castings were then mounted on the full-scale model in the F1 tunnel, and aerodynamic performance measurements were made using model surface pressure taps, the facility force balance system, and a large wake rake designed specifically for these tests. Tests were run over a range of Reynolds and Mach numbers. For each run, the model was rotated over a range of angles-of-attack that included airfoil stall. The benchmark data collected during these campaigns were, and continue to be, used for various purposes. The full-scale data form a unique, ice-accretion and associated aerodynamic performance dataset that can be used as a reference when addressing concerns regarding the use of subscale ice-accretion data to assess full-scale icing effects. Further, the data may be used in the development or enhancement of both ice-accretion prediction codes and computational fluid dynamic codes when applied to study the effects of icing. Finally, as was done in the wider study, the data may be used to help determine the level of geometric fidelity needed for artificial ice used to assess aerodynamic degradation due to aircraft icing. The structured, multifaceted approach used in this research effort provides a unique perspective on the aerodynamic effects of aircraft icing. The data presented in this report are available in electronic form upon formal approval by proper NASA and ONERA authorities.

Description: An approximately 6-percent scale model of the NASA Second-Generation Large Civil Tiltrotor (LCTR2) Aircraft was tested in the U.S. Army 7- by 10-Foot Wind Tunnel at NASA Ames Research Center January 4 to April 19, 2012, and September 18 to November 1, 2013. The full model was tested, along with modified versions in order to determine the effects of the wing tip extensions and nacelles; the wing was also tested separately in the various configurations. In both cases, the wing and nacelles used were adopted from the U.S. Army High Efficiency Tilt Rotor (HETR) aircraft, in order to limit the cost of the experiment. The full airframe was tested in high-speed cruise and low-speed hover flight conditions, while the wing was tested only in cruise conditions, with Reynolds numbers ranging from 0 to 1.4 million. In all cases, the external scale system of the wind tunnel was used to collect data. Both models were mounted to the scale using two support struts attached underneath the wing; the full airframe model also used a third strut attached at the tail. The collected data provides insight into the performance of the preliminary design of the LCTR2 and will be used for computational fluid dynamics (CFD) validation and the development of flight dynamics simulation models.

Description: In the interest of improving the predictability of high-lift systems at maximum lift conditions, a series of fundamental experiments were conducted to study the effects of adverse pressure gradient on a wake flow. Mean and fluctuating velocities were measured with a two-component laser-Doppler velocimeter. Data were obtained for several cases of adverse pressure gradient, producing flows ranging from no reversed flow to massively reversed flow. While the turbulent Reynolds stresses increase with increasing size of the reversed flow region, the gradient of Reynolds stress does not. Computations using various turbulence models were unable to reproduce the reversed flow.

Description: Predictions for Reynolds-stress and triple product turbulence models are compared for flows with significant rotational effects. Driver spinning cylinder flowfield and Zaets rotating pipe case are to be investigated at a minimum.

Description: Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, the need exists to understand the limitations of ground based testing by augmenting the analysis of experimental test results with Computational Fluid Dynamics (CFD) modeling. The immediate objective of the present research is to develop an XV-15 Tilt Rotor Research Aircraft rotor model for investigation of wind tunnel wall interference. The predicted performance of the XV-15 during various flight modes is compared to theoretical and experimental data. This research is performed to support wind tunnel tests scheduled for 2016. A mid-fidelity RANS solver, RotCFD, is used with an unsteady, incompressible flow model and a realizable k- turbulence model. The rotor is modeled using an actuator disk model or blade element model with a momentum source approach. In RotCFD the setup, grid generation and running of cases is faster than many CFD codes which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. Being able to accurately predict unsteady rotorcraft performance on desktop-class computers provides a quicker analysis of highly complex flows during the initial design phase.

Description: No abstract available

Description: No abstract available

Description: This paper is concerned with the high Reynolds number flow over a spanwise periodic array of roughness elements with inter-element spacing of the order of the local boundary-layer thickness. While earlier work by Goldstein, Sescu, Duck and Choudhari (2010) and Goldstein, Sescu, Duck and Choudhari (2011) was mainly concerned with smaller roughness heights that produced relatively weak distortions of the downstream flow, the focus here is on extending the analysis to larger roughness heights and streamwise elongated planform shapes that together produce a qualitatively different, nonlinear behavior of the downstream wakes. The roughness scale flow now has a novel triple-deck structure that is somewhat different from related studies that have previously appeared in the literature. The resulting flow is formally nonlinear in the intermediate wake region, where the streamwise distance is large compared to the roughness dimensions but small compared to the downstream distance from the leading edge, as well as in the far wake region where the streamwise length scale is of the order of the downstream distance from the leading edge. In contrast, the flow perturbations in both of these wake regions were strictly linear in the earlier work by Goldstein et al (2010, 2011). This is an important difference because the nonlinear wake flow in the present case provides an appropriate basic state for studying the secondary instability and eventual breakdown into turbulence.

Description: NASA conducted a winter 2015 field campaign using weather balloons at the NASA Glenn Research Center to generate a validation database for the NASA Icing Remote Sensing System. The weather balloons carried a specialized, disposable, vibrating-wire sensor to determine supercooled liquid water content aloft. Significant progress has been made to calibrate and characterize these sensors. Calibration testing of the vibrating-wire sensors was carried out in a specially developed, low-speed, icing wind tunnel, and the results were analyzed. The sensor ice accretion behavior was also documented and analyzed. Finally, post-campaign evaluation of the balloon soundings revealed a gradual drift in the sensor data with increasing altitude. This behavior was analyzed and a method to correct for the drift in the data was developed.

Description: An In-Situ Load System for calibrating and validating aerodynamic properties of scaled aircraft in ground-based aerospace testing applications includes an assembly having upper and lower components that are pivotably interconnected. A test weight can be connected to the lower component to apply a known force to a force balance. The orientation of the force balance can be varied, and the measured forces from the force balance can be compared to applied loads at various orientations to thereby develop calibration factors.

Description: On October 23, 2015, the Dawn spacecraft left the High Altitude Mapping Orbit (HAMO) around Ceres and began its final decent to the Low Altitude Mapping Orbit (LAMO), arriving on December 15. The transfer between the two science orbits, a tight spiraling trajectory with over 100 revolutions, required the operations team to perform weekly maneuver designs for a period of 50 days. While the first six weeks of the transfer executed as planned, unexpectedly the spacecraft incurred a multi-sigma delivery error to the final science orbit that was subsequently clean-up at the first orbit maintenance maneuver. In this paper we discuss the design architecture for the transfer in detail, including challenges the team faced in flying the transfer and lessons learned.

Description: New guidance of acceptable means of compliance with the super-cooled large drops (SLD) conditions has been issued by the U.S. Department of Transportation's Federal Aviation Administration (FAA) in its Advisory Circular AC 25-28 in November 2014. The Part 25, Appendix O is developed to define a representative icing environment for super-cooled large drops. Super-cooled large drops, which include freezing drizzle and freezing rain conditions, are not included in Appendix C. This paper reports results from recent glaze icing scaling tests conducted in NASA Glenn Icing Research Tunnel (IRT) to evaluate how well the scaling methods recommended for Appendix C conditions might apply to SLD conditions. The models were straight NACA 0012 wing sections. The reference model had a chord of 72 in. and the scale model had a chord of 21 in. Reference tests were run with airspeeds of 100 and 130.3 kn and with MVD's of 85 and 170 micron. Two scaling methods were considered. One was based on the modified Ruff method with scale velocity found by matching the Weber number WeL. The other was proposed and developed by Feo specifically for strong glaze icing conditions, in which the scale liquid water content and velocity were found by matching reference and scale values of the nondimensional water-film thickness expression and the film Weber number Wef. All tests were conducted at 0 deg AOA. Results will be presented for stagnation freezing fractions of 0.2 and 0.3. For nondimensional reference and scale ice shape comparison, a new post-scanning ice shape digitization procedure was developed for extracting 2-D ice shape profiles at any selected span-wise location from the high fidelity 3-D scanned ice shapes obtained in the IRT.

Description: An overview of recent progress regarding the computational aeroelastic and aeroservoelastic (ASE) analyses of a low-boom supersonic configuration is presented. The overview includes details of the computational models developed to date with a focus on unstructured CFD grids, computational aeroelastic analyses, sonic boom propagation studies that include static aeroelastic effects, and gust loads analyses. In addition, flutter boundaries using aeroelastic Reduced-Order Models (ROMs) are presented at various Mach numbers of interest. Details regarding a collaboration with the Royal Institute of Technology (KTH, Stockholm, Sweden) to design, fabricate, and test a full-span aeroelastic wind-tunnel model are also presented.

Description: Direct numerical simulations (DNS) are used to examine the turbulence statistics and the radiation field generated by a high-speed turbulent boundary layer with a nominal freestream Mach number of 14 and wall temperature of 0:18 times the recovery temperature. The flow conditions fall within the range of nozzle exit conditions of the Arnold Engineering Development Center (AEDC) Hypervelocity Tunnel No. 9 facility. The streamwise domain size is approximately 200 times the boundary-layer thickness at the inlet, with a useful range of Reynolds number corresponding to Re 450 650. Consistent with previous studies of turbulent boundary layer at high Mach numbers, the weak compressibility hypothesis for turbulent boundary layers remains applicable under this flow condition and the computational results confirm the validity of both the van Driest transformation and Morkovin's scaling. The Reynolds analogy is valid at the surface; the RMS of fluctuations in the surface pressure, wall shear stress, and heat flux is 24%, 53%, and 67% of the surface mean, respectively. The magnitude and dominant frequency of pressure fluctuations are found to vary dramatically within the inner layer (z/delta 0.〈 or approx. 0.08 or z+ 〈 or approx. 50). The peak of the pre-multiplied frequency spectrum of the pressure fluctuation is f(delta)/U(sub infinity) approx. 2.1 at the surface and shifts to a lower frequency of f(delta)/U(sub infinity) approx. 0.7 in the free stream where the pressure signal is predominantly acoustic. The dominant frequency of the pressure spectrum shows a significant dependence on the freestream Mach number both at the wall and in the free stream.

Description: This paper presents a status report on the collaboration between the Royal Institute of Technology (KTH) in Sweden and the NASA Langley Research Center regarding the design, fabrication, modeling, and testing of a full-span lighter configuration in the Transonic Dynamics Tunnel (TDT). The goal of the test is to acquire transonic limit-cycle- oscillation (LCO) data, including accelerations, strains, and unsteady pressures. Finite element models (FEMs) and aerodynamic models are presented and discussed along with results obtained to date.

Description: The NASA Advanced Air Vehicles Program, Commercial Supersonics Technology Project seeks to advance tools and techniques to make over-land supersonic flight feasible. In this study, preliminary computational results are presented for future tests in the NASA Ames 9 foot x 7 foot supersonic wind tunnel to be conducted in early 2016. Shock-plume interactions and their effect on pressure signature are examined for six model geometries. Near- field pressure signatures are assessed using the CFD code USM3D to model the proposed test geometries in free-air. Additionally, results obtained using the commercial grid generation software Pointwise Reigistered Trademark are compared to results using VGRID, the NASA Langley Research Center in-house mesh generation program.

Description: Gradient-based sensitivity analysis has proven to be an enabling technology for many applications, including design of aerospace vehicles. However, conventional sensitivity analysis methods break down when applied to long-time averages of chaotic systems. This breakdown is a serious limitation because many aerospace applications involve physical phenomena that exhibit chaotic dynamics, most notably high-resolution large-eddy and direct numerical simulations of turbulent aerodynamic flows. A recently proposed methodology, Least Squares Shadowing (LSS), avoids this breakdown and advances the state of the art in sensitivity analysis for chaotic flows. The first application of LSS to a chaotic flow simulated with a large-scale computational fluid dynamics solver is presented. The LSS sensitivity computed for this chaotic flow is verified and shown to be accurate, but the computational cost of the current LSS implementation is high.

Description: This study extends an existing semi-empirical approach to high-lift analysis by examining its effectiveness for use with a three-dimensional aerodynamic analysis method. The aircraft high-lift geometry is modeled in Vehicle Sketch Pad (OpenVSP) using a newly-developed set of techniques for building a three-dimensional model of the high-lift geometry, and for controlling flap deflections using scripted parameter linking. Analysis of the low-speed aerodynamics is performed in FlightStream, a novel surface-vorticity solver that is expected to be substantially more robust and stable compared to pressure-based potential-flow solvers and less sensitive to surface perturbations. The calculated lift curve and drag polar are modified by an empirical lift-effectiveness factor that takes into account the effects of viscosity that are not captured in the potential-flow solution. Analysis results are validated against wind-tunnel data for The Energy-Efficient Transport AR12 low-speed wind-tunnel model, a 12-foot, full-span aircraft configuration with a supercritical wing, full-span slats, and part-span double-slotted flaps.

Description: A computational study was performed for a Hybrid Wing Body configuration that was focused at transonic cruise performance conditions. In the absence of experimental data, two fully independent computational fluid dynamics analyses were conducted to add confidence to the estimated transonic performance predictions. The primary analysis was performed by Boeing with the structured overset-mesh code OVERFLOW. The secondary analysis was performed by NASA Langley Research Center with the unstructured-mesh code USM3D. Both analyses were performed at full-scale flight conditions and included three configurations customary to drag buildup and interference analysis: a powered complete configuration, the configuration with the nacelle/pylon removed, and the powered nacelle in isolation. The results in this paper are focused primarily on transonic performance up to cruise and through drag rise. Comparisons between the CFD results were very good despite some minor geometric differences in the two analyses.

Description: No abstract available

Description: A computational design and analysis methodology is being developed to design a vehicle that can support significant regions of natural laminar flow (NLF) at supersonic flight conditions. The methodology is built in the CDISC design module to be used in this paper with the flow solvers Cart3D and USM3D, and the transition prediction modules BLSTA3D and LASTRAC. The NLF design technique prescribes a target pressure distribution for an existing geometry based on relationships between modal instability wave growth and pressure gradients. The modal instability wave growths (both on- and off-axes crossflow and Tollmien-Schlichting) are balanced to produce a pressure distribution that will have a theoretical maximum NLF region for a given streamwise wing station. An example application is presented showing the methodology on a generic supersonic transport wingbody configuration. The configuration has been successfully redesigned to support significant regions of NLF (approximately 40% of the wing upper surface by surface area). Computational analysis predicts NLF with transition Reynolds numbers (ReT) as high as 36 million with 72 degrees of leading-edge sweep (LE), significantly expanding the current boundary of ReT - LE combinations for NLF. This NLF geometry provides a total drag savings of 4.3 counts compared to the baseline wing-body configuration (approximately 5% of total drag). Off-design evaluations at near-cruise and low-speed, high-lift conditions are discussed, as well as attachment line contamination/transition concerns. This computational NLF design effort is a part of an ongoing cooperative agreement between NASA and JAXA researchers.

Description: Measurement systems are typically calibrated based on standard practices established by a metrology standards laboratory, for example the National Institute for Standards and Technology (NIST), or dictated by an organization's metrology manual. Therefore, the calibration is designed and executed according to an established procedure. However, for many aerodynamic research measurement systems a universally accepted standard, traceable approach does not exist. Therefore, a strategy for how to develop a calibration protocol is left to the developer or user to define based on experience and recommended practice in their respective industry. Wind tunnel balances are one such measurement system. Many different calibration systems, load schedules and procedures have been developed for balances with little consensus on a recommended approach. Especially lacking is guidance the number of calibration data points needed. Regrettably, the number of data points tends to be correlated with the perceived quality of the calibration. Often, the number of data points is associated with ones ability to generate the data rather than by a defined need in support of measurement objectives. Hence the title of the paper was conceived to challenge recent observations in the wind tunnel balance community that shows an ever increasing desire for more data points per calibration absent of guidance to determine when there are enough. This paper presents fundamental concepts and theory to aid in the development of calibration procedures for wind tunnel balances and provides a framework that is generally applicable to the characterization and calibration of other measurement systems. Questions that need to be answered are for example: What constitutes an adequate calibration? How much data are needed in the calibration? How good is the calibration? This paper will assist a practitioner in answering these questions by presenting an underlying theory on how to evaluate a calibration based on objective measures. This will enable the developer and user to design calibrations with quantified performance in terms of their capability to meet the user's objectives and a basis for comparing existing calibrations that may have been developed in an ad-hoc manner.

Description: No abstract available

Description: No abstract available

Description: The Exo-Brake is a simple, non-propulsive means of de-orbiting small payloads from orbital platforms such as the International Space Station (ISS). Two de-orbiting experiments with fixed surface area Exo-Brakes have been successfully conducted in the last two years on the TechEdSat-3 and -4 nano-satellite missions. The development of the free molecular flow aerodynamic data-base is presented in terms of angle of attack, projected front surface area variation, and altitude. Altitudes are considered ranging from the 400km ISS jettison altitude to 90km. Trajectory tools are then used to predict de-orbit/entry corridors with the inclusion of the key atmospheric and geomagnetic uncertainties. Control system strategies are discussed which will be applied to the next two planned TechEdSat-5 and -6 nano-satellite missions - thus increasing the targeting accuracy at the Von Karman altitude through the proposed drag modulation technique.

Description: An overview of recent applications of the FUN3D CFD code to computational aeroelastic, sonic boom, and aeropropulsoservoelasticity (APSE) analyses of a low-boom supersonic configuration is presented. The overview includes details of the computational models developed including multiple unstructured CFD grids suitable for aeroelastic and sonic boom analyses. In addition, aeroelastic Reduced-Order Models (ROMs) are generated and used to rapidly compute the aeroelastic response and utter boundaries at multiple flight conditions.

Description: No abstract available

Description: The reduction of the aerodynamic load that acts on a generic rotorcraft fuselage by the application of active flow control was investigated in a wind tunnel test conducted on an approximately 1/3-scale powered rotorcraft model simulating forward flight. The aerodynamic mechanisms that make these reductions, in both the drag and the download, possible were examined in detail through the use of the measured surface pressure distribution on the fuselage, velocity field measurements made in the wake directly behind the ramp of the fuselage and computational simulations. The fuselage tested was the ROBIN-mod7, which was equipped with a series of eight slots located on the ramp section through which flow control excitation was introduced. These slots were arranged in a U-shaped pattern located slightly downstream of the baseline separation line and parallel to it. The flow control excitation took the form of either synthetic jets, also known as zero-net-mass-flux blowing, and steady blowing. The same set of slots were used for both types of excitation. The differences between the two excitation types and between flow control excitation from different combinations of slots were examined. The flow control is shown to alter the size of the wake and its trajectory relative to the ramp and the tailboom and it is these changes to the wake that result in a reduction in the aerodynamic load.

Description: Wind tunnel tests of a 5.75% scale model of the Boeing Hybrid Wing Body (HWB) configuration were conducted in the NASA Langley Research Center (LaRC) 14'x22' and NASA Ames Research Center (ARC) 40'x80' low speed wind tunnels as part of the NASA Environmentally Responsible Aviation (ERA) Project. Computational fluid dynamics (CFD) simulations of the flow-through nacelle (FTN) configuration of this model were performed before and after the testing. This paper presents a summary of the experimental and CFD results for the model in the cruise and landing configurations.

Description: In the field of computational fluid dynamics, the Navier-Stokes equations are often solved using an unstructuredgrid approach to accommodate geometric complexity. Implicit solution methodologies for such spatial discretizations generally require frequent solution of large tightly-coupled systems of block-sparse linear equations. The multicolor point-implicit solver used in the current work typically requires a significant fraction of the overall application run time. In this work, an efficient implementation of the solver for graphics processing units is proposed. Several factors present unique challenges to achieving an efficient implementation in this environment. These include the variable amount of parallelism available in different kernel calls, indirect memory access patterns, low arithmetic intensity, and the requirement to support variable block sizes. In this work, the solver is reformulated to use standard sparse and dense Basic Linear Algebra Subprograms (BLAS) functions. However, numerical experiments show that the performance of the BLAS functions available in existing CUDA libraries is suboptimal for matrices representative of those encountered in actual simulations. Instead, optimized versions of these functions are developed. Depending on block size, the new implementations show performance gains of up to 7x over the existing CUDA library functions.

Description: A new Isokinetic Total Water Content Evaporator (IKP2) was downsized from a prototype instrument, specifically to make airborne measurements of hydrometeor total water content (TWC) in deep tropical convective clouds to assess the new ice crystal Appendix D icing envelope. The probe underwent numerous laboratory and wind tunnel investigations to ensure reliable operation under the difficult high altitude/speed/TWC conditions under which other TWC instruments have been known to either fail, or have unknown performance characteristics and the results are presented in a companion paper (Ref. 1). This paper presents the equations used to determine the total water content (TWC) of the sampled atmosphere from the values measured by the IKP2 or necessary ancillary data from other instruments. The uncertainty in the final TWC is determined by propagating the uncertainty in the measured values through the calculations to the final result. Two techniques were used and the results compared. The first is a typical analytical method of propagating uncertainty and the second performs a Monte Carlo simulation. The results are very similar with differences that are insignificant for practical purposes. The uncertainty is between 2 and 3 percent at most practical operating conditions. The capture efficiency of the IKP2 was also examined based on a computational fluid dynamic simulation of the original IKP and scaled down to the IKP2. Particles above 24 micrometers were found to have a capture efficiency greater than 99 percent at all operating conditions.

Description: NASA conducted a winter 2015 field campaign using weather balloons at the NASA Glenn Research Center to generate a validation database for the NASA Icing Remote Sensing System. The weather balloons carried a specialized, disposable, vibrating-wire sensor to determine supercooled liquid water content aloft. Significant progress has been made to calibrate and characterize these sensors. Calibration testing of the vibrating-wire sensors was carried out in a specially developed, low-speed, icing wind tunnel, and the results were analyzed. The sensor ice accretion behavior was also documented and analyzed. Finally, post-campaign evaluation of the balloon soundings revealed a gradual drift in the sensor data with increasing altitude. This behavior was analyzed and a method to correct for the drift in the data was developed.

Description: The second Cranked-Arrow Wing Aerodynamics Project, International, coordinated project has been underway to improve high-fidelity computational-fluid-dynamics predictions of slender airframe aerodynamics. The work is focused on two flow conditions and leverages a unique flight data set obtained with the F-16XL aircraft for comparison and validation. These conditions, a low-speed high-angle-of-attack case and a transonic low-angle-of-attack case, were selected from a prior prediction campaign wherein the computational fluid dynamics failed to provide acceptable results. In revisiting these two cases, approaches for improved results include better, denser grids using more grid adaptation to local flow features as well as unsteady higher-fidelity physical modeling like hybrid Reynolds-averaged Navier-Stokes/unsteady Reynolds-averaged Navier-Stokes/large-eddy simulation methods. The work embodies predictions from multiple numerical formulations that are contributed from multiple organizations where some authors investigate other possible factors that could explain the discrepancies in agreement (e.g., effects due to deflected control surfaces during the flight tests as well as static aeroelastic deflection of the outer wing). This paper presents the synthesis of all the results and findings and draws some conclusions that lead to an improved understanding of the underlying flow physics, finally making the connections between the physics and aircraft features.

Description: A data acquisition system upgrade project, known as AB-DAS, is underway at the NASA Langley Transonic Dynamics Tunnel. AB-DAS will soon serve as the primary data system and will substantially increase the scan-rate capabilities and analog channel count while maintaining other unique aeroelastic and dynamic test capabilities required of the facility. AB-DAS is configurable, adaptable, and enables buffet and aeroacoustic tests by synchronously scanning all analog channels and recording the high scan-rate time history values for each data quantity. AB-DAS is currently available for use as a stand-alone data system with limited capabilities while development continues. This paper describes AB-DAS, the design methodology, and the current features and capabilities. It also outlines the future work and projected capabilities following completion of the data system upgrade project.

Description: The linear form of parabolized linear stability equations (PSE) is used in a variational approach to extend the previous body of results for the optimal, non-modal disturbance growth in boundary layer flows. This methodology includes the non-parallel effects associated with the spatial development of boundary layer flows. As noted in literature, the optimal initial disturbances correspond to steady counter-rotating stream-wise vortices, which subsequently lead to the formation of stream-wise-elongated structures, i.e., streaks, via a lift-up effect. The parameter space for optimal growth is extended to the hypersonic Mach number regime without any high enthalpy effects, and the effect of wall cooling is studied with particular emphasis on the role of the initial disturbance location and the value of the span-wise wavenumber that leads to the maximum energy growth up to a specified location. Unlike previous predictions that used a basic state obtained from a self-similar solution to the boundary layer equations, mean flow solutions based on the full Navier-Stokes (NS) equations are used in select cases to help account for the viscous-inviscid interaction near the leading edge of the plate and also for the weak shock wave emanating from that region. These differences in the base flow lead to an increasing reduction with Mach number in the magnitude of optimal growth relative to the predictions based on self-similar mean-flow approximation. Finally, the maximum optimal energy gain for the favorable pressure gradient boundary layer near a planar stagnation point is found to be substantially weaker than that in a zero pressure gradient Blasius boundary layer.

Description: The NASA Environmentally Responsible Aircraft Project (ERA) was a ve year project broken into two phases. In phase II, high N+2 Technical Readiness Level demonstrations were grouped into Integrated Technology Demonstrations (ITD). This paper describes the work done on ITD-51A: the Vehicle Systems Integration, Engine Airframe Integration Demonstration. Refinement of a Hybrid Wing Body (HWB) aircraft from the possible candidates developed in ERA Phase I was continued. Scaled powered, and unpowered wind- tunnel testing, with and without acoustics, in the NASA LARC 14- by 22-foot Subsonic Tunnel, the NASA ARC Unitary Plan Wind Tunnel, and the 40- by 80-foot test section of the National Full-Scale Aerodynamics Complex (NFAC) in conjunction with very closely coupled Computational Fluid Dynamics was used to demonstrate the fuel burn and acoustic milestone targets of the ERA Project.

Description: As part of the NASA Environmentally Responsible Aircraft project, an ultra high bypass ratio engine integration on a hybrid wing body demonstration was planned. The goal was to include engine and airframe integration concepts that reduced fuel consumption by at least 50% while still reducing noise 42 db cumulative on the ground. Since the engines would be mounted on the upper surface of the aft body of the aircraft, the inlets may be susceptible to vortex ingestion from the wing leading edge at high angles of attack and sideslip, and separated wing/body flow. Consequently, experimental and computational studies were conducted to collect flow surveys useful for characterizing engine operability. The wind tunnel tests were conducted at two NASA facilities, the 14- by 22-foot at NASA Langley and the 40- by 80-foot at NASA Ames Research Center. The test results included in this paper show that the distortion and pressure recovery levels were acceptable for engine operability. The CFD studies conducted to compare to experimental data showed excellent agreement for the angle of attacks examined, although failed to match the low speed experimental data at high sideslip angles.

Description: NASA has been working toward designing and conducting a juncture flow experiment on a wing-body aircraft configuration. The experiment is planned to provide validation-quality data for CFD that focuses on the onset and progression of a separation bubble near the wing-body juncture trailing edge region. This paper describes the goals and purpose of the experiment. Although currently considered unreliable, preliminary CFD analyses of several different configurations are shown. These configurations have been subsequently tested in a series of "risk-reduction" wind tunnel tests, in order to help down-select to a final configuration that will attain the desired flow behavior. The risk-reduction testing at the higher Reynolds number has not yet been completed (at the time of this writing), but some results from one of the low-Reynolds-number experiments are shown.

Description: An overview of twenty years of adjoint-based aerodynamic design research at NASA Langley Research Center is presented. Adjoint-based algorithms provide a powerful tool for efficient sensitivity analysis of complex large-scale computational fluid dynamics (CFD) simulations. Unlike alternative approaches for which computational expense generally scales with the number of design parameters, adjoint techniques yield sensitivity derivatives of a simulation output with respect to all input parameters at the cost of a single additional simulation. With modern large-scale CFD applications often requiring millions of compute hours for a single analysis, the efficiency afforded by adjoint methods is critical in realizing a computationally tractable design optimization capability for such applications.

Description: Airframe noise corresponds to the acoustic radiation due to turbulent flow in the vicinity of airframe components such as high-lift devices and landing gears. Since 2010, the American Institute of Aeronautics and Astronautics has organized an ongoing series of workshops devoted to Benchmark Problems for Airframe Noise Computations (BANC). The BANC workshops are aimed at enabling a systematic progress in the understanding and high-fidelity predictions of airframe noise via collaborative investigations that integrate computational fluid dynamics, computational aeroacoustics, and in depth measurements targeting a selected set of canonical yet realistic configurations that advance the current state-of-the-art in multiple respects. Unique features of the BANC Workshops include: intrinsically multi-disciplinary focus involving both fluid dynamics and aeroacoustics, holistic rather than predictive emphasis, concurrent, long term evolution of experiments and simulations with a powerful interplay between the two, and strongly integrative nature by virtue of multi-team, multi-facility, multiple-entry measurements. This paper illustrates these features in the context of the BANC problem categories and outlines some of the challenges involved and how they were addressed. A brief summary of the BANC effort, including its technical objectives, strategy, and selective outcomes thus far is also included.

Description: Computational analysis of experimental aircraft prior to test ights can be a valuable tool to estimate ight characteristics and determine areas of elevated caution. It can also provide feedback to software and model developers as to the accuracy of models used when the aircraft is ultimately own. This paper describes the aerodynamic analysis and characterisation of an experimental tilt-wing aircraft with a unique design. The paper covers what analysis is performed as well as results of these aircraft characterisations. Through this analysis a database le is created for use with NASA Design and Analysis of Rotorcraft (NDARC) tool.

Description: NASA conducted a winter 2015 field campaign at the NASA Glenn Research Center intended to generate a validation database for the NASA Icing Remote Sensing System. The weather balloons carried a specialized, disposable sensor designed to determine supercooled liquid water content aloft. Significant progress has been made to calibrate and characterize these specialized sensors. Calibration testing of these sensors was carried out in a specially developed, low-speed, icing wind tunnel. The sensor icing behavior was documented and analyzed. Finally, post-campaign evaluation of the balloon soundings revealed a gradual drift in the sensor data with increasing altitude. The behavior was analyzed and a method to account for the drift was developed.

Description: Experimental techniques to measure rotorcraft aerodynamic performance are widely used. However, most of them are either unable to capture interference effects from bodies, or require an extremely large computational budget. The objective of the present research is to develop an XV-15 Tiltrotor Research Aircraft rotor model for investigation of wind tunnel wall interference using a novel Computational Fluid Dynamics (CFD) solver for rotorcraft, RotCFD. In RotCFD, a mid-fidelity Unsteady Reynolds Averaged Navier-Stokes (URANS) solver is used with an incompressible flow model and a realizable k- turbulence model. The rotor is, however, not modeled using a computationally expensive, unsteady viscous body-fitted grid, but is instead modeled using a blade-element model (BEM) with a momentum source approach. Various flight modes of the XV-15 isolated rotor, including hover, tilt, and airplane mode, have been simulated and correlated to existing experimental and theoretical data. The rotor model is subsequently used for wind tunnel wall interference simulations in the National Full-Scale Aerodynamics Complex (NFAC) at Ames Research Center in California. The results from the validation of the isolated rotor performance showed good correlation with experimental and theoretical data. The results were on par with known theoretical analyses. In RotCFD the setup, grid generation, and running of cases is faster than many CFD codes, which makes it a useful engineering tool. Performance predictions need not be as accurate as high-fidelity CFD codes, as long as wall effects can be properly simulated. For both test sections of the NFAC wall, interference was examined by simulating the XV-15 rotor in the test section of the wind tunnel and with an identical grid but extended boundaries in free field. Both cases were also examined with an isolated rotor or with the rotor mounted on the modeled geometry of the Tiltrotor Test Rig (TTR). A "quasi linear trim" was used to trim the thrust for the rotor to compare the power as a unique variable. Power differences between free field and wind tunnel cases were found from -7 to 0 percent in the 80- by 120-Foot Wind Tunnel and -1.6 to 4.8 percent in the 40- by 80-Foot Wind Tunnel, depending on the TTR orientation, tunnel velocity, and blade setting. The TTR will be used in 2016 to test the Bell 609 rotor in a similar fashion to the research in this report.

Description: A study was undertaken to investigate the measurement of wing deformation and internal loads using measured strain data. Future aerospace vehicle research depends on the ability to accurately measure the deformation and internal loads during ground testing and in flight. The approach uses the inverse Finite Element Method (iFEM). The iFEM is a robust, computationally efficient method that is well suited for real-time measurement of real-time structural deformation and loads. The method has been validated in previous work, but has yet to be applied to a large-scale test article. This work is in preparation for an upcoming loads test of a half-span test wing in the Flight Loads Laboratory at the National Aeronautics and Space Administration Armstrong Flight Research Center (Edwards, California). The method has been implemented into an efficient MATLAB (The MathWorks, Inc., Natick, Massachusetts) code for testing different sensor configurations. This report discusses formulation and implementation along with the preliminary results from a representative aerospace structure. The end goal is to investigate the modeling and sensor placement approach so that the best practices can be applied to future aerospace projects.

Description: The near wake of a flat plate with circular and elliptic trailing edges is investigated with data from direct numerical simulations. The plate length and thickness are the same in both cases. The separating boundary layers are turbulent and statistically identical. Therefore the wake is symmetric in the two cases. The emphasis in this study is on a comparison of the wake-distributions of velocity components, normal intensity and fluctuating shear stress obtained in the two cases.

Description: A 0.656-scale V-22 proprotor, the Joint Vertical Experimental (JVX) rotor, was tested at the NASA Ames Research Center in both hover and airplane-mode (high-speed axial flow) flight conditions, up to an advance ratio of 0.562 (231 knots). This paper examines the two principal data sets generated by those tests, and includes investigations of hub spinner tares, torque/thrust measurement interactions, tunnel blockage effects, and other phenomena suspected of causing erroneous measurements or predictions. Uncertainties in hover and high-speed data are characterized. The results are reported here to provide guidance for future wind tunnel tests, data processing, and data analysis.

Description: This presentation reports results from recent icing scaling tests in NASA Glenn Icing Research Tunnel (IRT) to evaluate how well the scaling method recommended for Appendix C conditions might apply to SLD conditions.

Description: The greatest efficiency for a lifting surface at supersonic speeds, according to the theoretical considerations of reference 1, can be attained if the leading edge is swept well behind the Mach cone and the highest aspect ratio which is structurally possible is employed. Such a wing, designed for a Mach number of 3.0, would have 80 deg. of sweepback. Aeroelastic effects have 〈 been shown 3 to be considerable for a wing with 60deg of sweepback and designed for a Mach number of 2.0. The wing shown was found theoretically to have considerable loss in maximum lift-drag ratio attributable to aeroelasticity. This wing has 12-per cent-thick Clark-Y airfoils normal to the wing leading edge. If it were of solid aluminum and flying at a dynamic pressure of 2,400 lbs./sq.ft. (flexibility parameter qb(exp. 4) /El(0) = 7.8), analysis indicates that the wing would deflect so as to reduce the maximum lift-drag ratio about 30 per cent.

Description: Measurements of average skin friction of the turbulent boundary layer have been made on a 15deg total included angle cone with foreign gas injection. Measurements of total skin-friction drag were obtained at free-stream Mach numbers of 0.3, 0.7, 3.5, and 4.7 and within a Reynolds number range from 0.9 x 10(exp 6) to 5.9 x 10(exp 6) with injection of helium, air, and Freon-12 (CCl2F2) through the porous wall. Substantial reductions in skin friction are realized with gas injection within the range of Mach numbers of this test. The relative reduction in skin friction is in accordance with theory-that is, the light gases are most effective when compared on a mass flow basis. There is a marked effect of Mach number on the reduction of average skin friction; this effect is not shown by the available theories. Limited transition location measurements indicate that the boundary layer does not fully trip with gas injection but that the transition point approaches a forward limit with increasing injection. The variation of the skin-friction coefficient, for the lower injection rates with natural transition, is dependent on the flow Reynolds number and type of injected gas; and at the high injection rates the skin friction is in fair agreement with the turbulent boundary layer results.

Description: The aerodynamic effects of fixing boundary-layer transition for a swept- and a triangular-wing configuration have been determined from tests of two small-scale wing-body models. The wings had an aspect ratio of 2.99 and 3-percent-thick biconvex sections. Lift, pitching-moment, and drag data were obtained at Mach numbers ranging from 0.60 to 1.40 for angles of attack between -2 deg and about 15 deg. The Reynolds number of the tests was generally 1.5 million; however, minimum drag measurements were made for both models over a range of Reynolds numbers from 1.0 million to about 3.0 or 4.0 million.

Description: A theoretical analysis indicates that, for rotors, ground effect decreases rapidly with increases in either height above the ground or forward speed. The decrease with height above the ground in forward flights is greater than that in hovering. The major part of the decrease in ground effect with forward speed occurs at speeds less than 1.5 times the hovering mean induced velocity. Consequently, the total induced velocity at the rotor center increases rather than decreases when a helicopter gathers speed at low height above the ground.

Description: An investigation has been made to study the effect of ground proximity on the aerodynamic characteristics of two jet vertical-take-off-and-landing airplane models in which the fuselage remains in a horizontal attitude for the take-off and landing. The first model (called the tilt-wing model) had a tilting wing-engine assembly which was set at 90 deg incidence for the take-off and landing. The second model, called the deflected-jet model) had a cascade of retractable turning vanes to deflect the exhaust of the horizontally mounted jet engines downward for vertical take-off and landing while the entire model remained in a horizontal attitude. With the models at various heights above the ground in the take-off and landing configuration, the lift, drag, and pitching moment were measured and tuft surveys were made to determine the flow field caused by the jet exhaust. The tilt-wing model experienced a loss of lift of less than 3 percent near the ground. The deflected-jet model, however, suffered losses in lift as high as 45 percent near the ground because of a low pressure region under the model caused by the entrainment of air by the jet exhaust as it spread out along the ground. This loss in lift for the deflected-jet configuration could probably be reduced to less than 5 percent by the use of a longer landing gear and a high wing location.

Description: For the test, the 12-inch-diameter "Vortex-Ring" parachute was towed behind a conical-nosed cylindrical body 2.25 inches in diameter. The tow-cable length was 24 inches, and was attached to the cylindrical body through a large swivel and to the parachute through a smaller swivel. The attachment between the large swivel an the cylindrical body failed after about 1 minute's operation. Mach number was approximately 2.2, dynamic pressure was approximately 150 pounds per square foot, and camera speed was approximately 3000 frames per second.

Description: An exploratory investigation has been made in the Langley 300 MPH 7 by 10 foot tunnel to study the low-speed static longitudinal and lateral stability characteristics of a reentry configuration having rigid retractable conical lifting surfaces that unfolded from the surface of a conical fuselage. The model also had curved tail surfaces that unfolded from a cylindrical aft section attached to the cone. Longitudinal tests were made through an angle-of-attack range from -4 deg to 90 deg and limited lateral tests were made through an angle-of-sideslip range from -12 deg to 32 deg at an angle of attack of 0 deg. The tail surface provided longitudinal trim to maximum lift and beyond and up to an angle of attack of 51 deg for a center-of-moment location of 42.9 percent mean aerodynamic chord. For this center-of-moment position the model had a static margin of 12 percent mean aerodynamic chord at the lower lift coefficients and was longitudinally stable up to a lift coefficient between 1.0 and 1.2. Neutral stability occurred from lift coefficient of 1.0 up to near maximum lift coefficient. The maximum value of trimmed lift-drag ratio was 4.85 at a lift coefficient of approximately 0.3 and a trimmed angle of attack of approximately 10 deg. The configuration was directionally stable throughout the test angle of sideslip range for an angle of attack of 0 deg.

Description: This report describes a technique which combines theory and experiments for determining relaxation times in gases. The technique is based on the measurement of shapes of the bow shock waves of low-fineness-ratio cones fired from high-velocity guns. The theory presented in the report provides a means by which shadowgraph data showing the bow waves can be analyzed so as to furnish effective relaxation times. Relaxation times in air were obtained by this technique and the results have been compared with values estimated from shock tube measurements in pure oxygen and nitrogen. The tests were made at velocities ranging from 4600 to 12,000 feet per second corresponding to equilibrium temperatures from 35900 R (19900 K) to 6200 R (34400 K), under which conditions, at all but the highest temperatures, the effective relaxation times were determined primarily by the relaxation time for oxygen and nitrogen vibrations.

Description: A wind-tunnel investigation has been made to study the static longitudinal and lateral stability characteristics of a simplified aerial vehicle supported by ducted fans that tilt relative to the airframe. The ducts were in a triangular arrangement with one duct in front and two at the rear in order to minimize the influence of the downwash of the front duct on the rear ducts. The results of the investigation were compared with those of a similar investigation for a tandem two-duct arrangement in which the ducts were fixed (rather than tiltable) relative to the airframe, since the three-duct configuration had been devised in an attempt to avoid some of the deficiencies of the tandem fixed-duct configuration. The results of the investigation indicated that the tilting-duct arrangement had less noseup pitching moment for a given forward speed than the tandem fixed-duct arrangement. The model had less angle-of-attack instability than the tandem fixed-duct arrangement. The model was directionally unstable but had a positive dihedral effect throughout the test speed range.

Description: An investigation was made at high subsonic speeds in the Langley high-speed 7- by 10-foot tunnel to determine the effect of end plates on the longitudinal aerodynamic characteristics of a sweptback wing-body combination with and without drooped chord-extensions. The wing had 45 deg sweepback of the quarter-chord line, an aspect ratio of 4, a taper ratio of 0.3, and NACA 65AO06 airfoil sections parallel to the plane of symmetry, and was mounted near the rear of a body of revolution having a fineness ratio of approximately 8. The results indicated that the addition of the end plates to either the wing with drooped chord-extensions or to the wing without drooped chord-extensions slightly increased the lift in the low angle-of-attack range but slightly decreased the lift at moderate and high angles of attack. The addition of the end plates to the wing without the chord-extensions caused a small increase in the maximum lift-drag ratio at Mach numbers below 0.65 and a slight decrease at the higher Mach numbers; however, for the addition of the end plates to the wing with the chord- extensions the maximum lift-drag ratio was slightly decreased below a Mach number of 0.88, while a slight increase occurred for the higher Mach numbers. The addition of the end plates to the wings with and without the chord-extensions caused the static longitudinal stability to increase considerably for all Mach numbers; however, only a slight reduction in the aerodynamic-center variation with Mach number was observed.

Description: The experimental wave drags of bodies and wing-body combinations over a wide range of Mach numbers are compared with the computed drags utilizing a 24-term Fourier series application of the supersonic area rule and with the results of equivalent-body tests. The results indicate that the equivalent-body technique provides a good method for predicting the wave drag of certain wing-body combinations at and below a Mach number of 1. At Mach numbers greater than 1, the equivalent-body wave drags can be misleading. The wave drags computed using the supersonic area rule are shown to be in best agreement with the experimental results for configurations employing the thinnest wings. The wave drags for the bodies of revolution presented in this report are predicted to a greater degree of accuracy by using the frontal projections of oblique areas than by using normal areas. A rapid method of computing wing area distributions and area-distribution slopes is given in an appendix.

Description: A flight investigation has been conducted to study the heat transfer to swept-wing leading edges. A rocket-powered model was used for the investigation and provided data for Mach number ranges of 1.78 to 2.99 and 2.50 to 4.05 with corresponding free-stream Reynolds number per foot ranges of 13.32 x 10(exp 6) to 19.90 x 10(exp 6) and 2.85 x 10(exp 6) to 4.55 x 10(exp 6). The leading edges employed were cylindrically blunted wedges ', three of which were swept 450 with leading-edge diameters of 1/4, 1/2, and 3/4 inch and one swept 36-750 with a leading-edge diameter of 1/2 inch. In the high Reynolds number range, measured values of heat transfer were found to be much higher than those predicted by laminar theory and at the larger values of leading-edge diameter were approaching the values predicted by turbulent theory. For the low Reynolds number range a comparison between measured and theoretical heat transfer showed that increasing the leading-edge diameter resulted in turbulent flow on the cylindrical portion of the leading edge.

Description: The shock-wave patterns of a complex configuration with cranked cruciform wings and a cone-cylinder body were examined to determine the interaction of the body bow wave with the flow field about the wing. Also of interest, was the interaction of the forward (760 sweptback) wing leading-edge wave with the rear (600 sweptback) wing leading-edge wave. The shadowgraph pictures of the model in free flight at a Mach number of 4.9, although not definitive, appear to indicate that the body bow wave crosses the outer wing panel after first being refracted either by the leading-edge wave of the 600 sweptback wing or by pressure fields in the flow crossing the wing.

Description: Experimental results are presented for an exploratory investigation of the effectiveness of interference between jet and afterbody in reducing the axial force on an afterbody with a neighboring jet. In addition to the interference axial force., measurements are presented of the interference normal force and the center of pressure of the interference normal force. The free-stream Mach number was 2.94, the jet-exit Mach number was 2.71, and the Reynolds number was 0.25 x 10, based on body diameter. The variables investigated include static-pressure ratio of the jet (up to 9), nacelle position relative to afterbody, angle of attack (-5 deg to 10 deg), and afterbody shape. Two families of afterbody shapes were tested. One family consisted of tangent-ogive bodies of revolution with varying length and base areas. The other family was formed by taking a planar slice off a circular cylinder with varying angle between the plane and cylinder. The trends with these variables are shown for conditions near maximum jet-afterbody interference. The interference axial forces are large and favorable. For several configurations the total afterbody axial force is reduced to zero by the interference.

Description: A wind-tunnel investigation was conducted to determine the effect of trailing-edge flaps with blowing-type boundary-layer control and leading-edge slats on the low-speed performance of a large-scale jet transport model with four engines and a 35 deg. sweptback wing of aspect ratio 7. Two spanwise extents and several deflections of the trailing-edge flap were tested. Results were obtained with a normal leading-edge and with full-span leading-edge slats. Three-component longitudinal force and moment data and boundary-layer-control flow requirements are presented. The test results are analyzed in terms of possible improvements in low-speed performance. The effect on performance of the source of boundary-layer-control air flow is considered in the analysis.

Description: This investigation is a continuation of the experimental and theoretical evaluation of the effects of wing plan-form variations on the aerodynamic performance characteristics of blended wing-body combinations. The present report compares previously tested straight-edged delta and arrow models which have leading-edge sweeps of 59.04 and 70-82 deg., respectively, with related models which have plan forms with curved leading and trailing edges designed to result in the same average sweeps in each case. All the models were symmetrical, without camber, and were generally similar having the same span, length, and aspect ratios. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local wing chords in a streamwise direction. The wing sections were computed by varying their shapes along with the body radii (blending process) to match the selected area distribution and the given plan form. The models were tested with transition fixed at Reynolds numbers of roughly 4,000,000 to 9,000,000, based on the mean aerodynamic chord of the wing. The characteristic effect of the wing curvature of the delta and arrow models was an increase at subsonic and transonic speeds in the lift-curve slopes which was partially reflected in increased maximum lift-drag ratios. Curved edges were not evaluated on a diamond plan form because a preliminary investigation indicated that the curvature considered would increase the supersonic zero-lift wave drag. However, after the test program was completed, a suitable modification for the diamond plan form was discovered. The analysis presented in the appendix indicates that large reductions in the zero-lift wave drag would be obtained at supersonic Mach numbers if the leading- and trailing-edge sweeps are made to differ by indenting the trailing edge and extending the root of the leading edge.

Description: An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the development of lift on a wing during a simulated constant-acceleration catapult take-off. The investigation included models of a two-dimensional wing, an unswept wing having an aspect ratio of 6, a 35 deg. swept wing having an aspect ratio of 3.05, and a 60 deg. delta wing having an aspect ratio of 2.31. All the wings investigated developed at least 90 percent of their steady-state lift in the first 7 chord lengths of travel. The development of lift was essentially independent of the acceleration when based on chord lengths traveled, and was in qualitative agreement with theory.

Description: Experimental research has been conducted on the effects of wall cooling, Mach number, and unit Reynolds number on the transition Reynolds number of cylindrical separated boundary layers on an ogive-cylinder model. Results were obtained from pressure and temperature measurements and shadowgraph observations. The maximum scope of measurements encompassed Mach numbers between 2.06 and 4.24, Reynolds numbers (based on length of separation) between 60,000 and 400,000, and ratios of wall temperature to adiabatic wall temperature between 0.35 and 1.0. Within the range of tile present tests, the transition Reynolds number was observed to decrease with increasing wall cooling, increase with increasing Mach number, and increase with increasing unit Reynolds number. The wall cooling effect was found to be four times as great when the attached boundary layer upstream of separation was cooled in conjunction with cooling of the separated boundary layer as when only the separated boundary layer was cooled. Wall cooling of both the attached and separated flow regions also caused, in some cases, reattachment in the otherwise separated region. Cavity resonance present in the separated region for some model configurations was accompanied by a large decrease in transition Reynolds number at the lower test Mach numbers.

Description: Large-scale wind-tunnel tests were made of a wingless vertical take-off and landing aircraft at zero sideslip to determine performance and longitudinal stability and control characteristics at airspeeds from 0 to 70 knots. Roll control and rudder effectiveness were also obtained. Limitations in the propulsion system restricted the lift for which level flight could be simulated to approximately 1500 pounds. Test variables with roll control and rudder undeflected were airspeed, vane setting, angle of attack, elevator deflection, and power. In most of the tests angle of attack, elevator, and power were varied individually while the other four parameters were held constant at previously determined values required for simulating trimmed level flight. The majority of the tests were made with power on and tail on at airspeeds between 20 and 70 knots. However, a limited number of data were obtained for the following conditions: (1) at zero velocity, horizontal tail on, power on; (2) at forward velocity, tail off and power on; and (3) at forward velocity, tail on, but with power off.

Description: Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.

Description: An investigation has been conducted in the Langley full-scale tunnel to determine the effects of a blowing boundary-layer-control lift-augmentation system on the aerodynamic characteristics of a large-scale model of a fighter-type airplane. The wing was unswept at the 70-percent- chord station, had an aspect ratio of 2.86, a taper ratio of 0.40, and 4-percent-thick biconvex airfoil sections parallel to the plane of symmetry. The tests were conducted over a range of angles of attack from approximately -4 deg to 23 deg for a Reynolds number of approximately 5.2 x 10(exp 6) which corresponds to a Mach number of 0.08. Blowing rates were normally restricted to values just sufficient to control air-flow separation. The results of this investigation showed that wing leading-edge blowing in combination with large values of wing leading-edge-flap deflection was a very effective leading-edge flow-control device for wings having highly loaded trailing-edge flaps. With leading-edge blowing there was no hysteresis of the lift, drag, and pitching-moment characteristics upon recovery from stall. End plates were found to improve the lift and drag characteristics of the test configuration in the moderate angle-of-attack range, and blockage to one-quarter of the blowing-slot area was not detrimental to the aerodynamic characteristics. Blowing boundary-layer control resulted in a considerably reduced landing speed and reduced landing and take-off distances. The ailerons were very effective lateral-control devices when used with blowing flaps.

Description: The problem of chordwise, or camber, divergence at transonic and supersonic speeds is treated with primary emphasis on slender delta wings having a cantilever support at the trailing edge. Experimental and analytical results are presented for four wing models having apex half-angles of 5 deg, 10 deg, 15 deg, and 20 deg. A Mach number range from 0.8 to 7.3 is covered. The analytical results include calculations based on small-aspect-ratio theory, lifting-surface theory, and strip theory. A closed-form solution of the equilibrium equation is given, which is based on low-aspect-ratio theory but which applies only to certain stiffness distributions. Also presented is an iterative procedure for use with other aerodynamic theories and with arbitrary stiffness distribution.

Description: A flutter analysis employing the kernel function for three- dimensional, subsonic, compressible flow is applied to a flutter-tested tail surface which has an aspect ratio of 3.5, a taper ratio of 0.15, and a leading-edge sweep of 30 deg. Theoretical and experimental results are compared at Mach numbers from 0.75 to 0.98. Good agreement between theoretical and experimental flutter dynamic pressures and frequencies is achieved at Mach numbers to 0.92. At Mach numbers from 0.92 to 0.98, however, a second solution to the flutter determinant results in a spurious theoretical flutter boundary which is at a much lower dynamic pressure and at a much higher frequency than the experimental boundary.

Description: Drag characteristics have been obtained for the X-15 airplane during unpowered flight. These data represent a Mach number range from about 0.7 to 3.1 and a Reynolds number range from 13.9 x 10(exp 6) to 28 x 10(exp 8), based on the mean aerodynamic chord. The full-scale data are compared with estimates compiled from several wind-tunnel facilities. The agreement between wind-tunnel and full-scale supersonic drag, uncorrected for Reynolds number effects, is reasonably close except at low supersonic Mach numbers where the flight values are significantly higher.

Description: The models had aspect-ratio-2 diamond, delta, and arrow wings with the leading edges swept 45.00 deg, 59.04 deg, and 70.82 deg, respectively. The wing sections were computed by varying the section shape along with the body radii (blending process) to match the prescribed area distribution and wing plan form. The wing sections had an average value of maximum thickness ratio of about 4 percent of the local chords in a streamwise direction. The models were tested with transition fixed at Reynolds numbers of about 4,000,000 to 9,000,0000, based on the mean aerodynamic chord of the wings. The effect of varying Reynolds number was checked at both subsonic and supersonic speeds. The diamond model was superior to the other plan forms at transonic speeds ((L/D)max = 11.00 to 9.52) because of its higher lift-curve slope and near optimum wave drag due to the blending process. For the wing thickness tested with the diamond model, the marked body and wing contouring required for transonic conditions resulted in a large wave-drag penalty at the higher supersonic Mach numbers where the leading and trailing edges of the wing were supersonic. Because of the low sweep of the trailing edge of the delta model, this configuration was less adaptable to the blending process. Removing a body bump prescribed by the Mach number 1.00 design resulted in a good supersonic design. This delta model with 10 percent less volume was superior to the other plan forms at Mach numbers of 1.55 to 2.35 ((L/D)max = 8.65 to 7.24), but it and the arrow model were equally good at Mach numbers of 2.50 to 3.50 ((L/D)max - 6.85 to O.39). At transonic speeds the arrow model was inferior because of the reduced lift-curve slope associated with its increased sweep and also because of the wing base drag. The wing base-drag coefficients of the arrow model based on the wing planform area decreased from a peak value of 0.0029 at Mach number 1.55 to 0.0003 at Mach number 3.50. Linear supersonic theory was satisfactory for predicting the aerodynamic trends at Mach numbers from 1.55 to 3.50 of lift-curve slope, wave drag, drag due to lift, aerodynamic-center location, and maximum lift-drag ratios for each of the models.

Description: An experimental investigation was performed at a Mach number of 3.0 to determine the friction and pressure drags of a pylon and a 20 deg- and a 40 deg-included-angle wedge diverter over a range of Reynolds number. The results indicated that the measured friction drag coefficients agreed reasonably with that predicted by flat-plate theory. The pressure drag coefficients of the 20 and 40 deg wedges agreed with those presented in the literature. The total drag coefficient of the pylon and the 20 deg wedge diverter was about 0.36, based on diverter frontal area, while the drag coefficient of the 40 deg wedge was about 0.47.

Description: An investigation has been made in the Langley 20-foot free-spinning tunnel to determine the erect and inverted spin and recovery characteristics of a 1/30-scale dynamic model of a twin-jet swept-wing fighter airplane. The model results indicate that the optimum erect spin recovery technique determined (simultaneous rudder reversal to full against the spin and aileron deflection to full with the spin) will provide satisfactory recovery from steep-type spins obtained on the airplane. It is considered that the air-plane will not readily enter flat-type spins, also indicated as possible by the model tests, but developed-spin conditions should be avoided in as much as the optimum recovery procedure may not provide satisfactory recovery if the airplane encounters a flat-type developed spin. Satisfactory recovery from inverted spins will be obtained on the airplane by neutralization of all controls. A 30-foot- diameter (laid-out-flat) stable tail parachute having a drag coefficient of 0.67 and a towline length of 27.5 feet will be satisfactory for emergency spin recovery.

Description: The intensity of shock-wave noise at the ground resulting from flights at Mach numbers to 2.0 and altitudes to 60,000 feet was measured. Meagurements near the ground track for flights of a supersonic fighter and one flight of a supersonic bomber are presented. Level cruising flight at an altitude of 60,000 feet and a Mach number of 2.0 produced sonic booms which were considered to be tolerable, and it is reasonable t o expect that cruising flight at higher altitudes will produce booms of tolerable intensity for airplanes of the size and weight of the test airplanes. The measured variation of sonic-boom intensity with altitude was in good agreement with the variation calculated by an equation given in NASA Technical Note D-48. The effect of Mach number on the ground overpressure is small between Mach numbers of 1.4 and 2.0, a result in agreement with the theory. No amplification of the shock-wave overpressures due to refraction effects was apparent near the cutoff Mach number. A method for estimating the effect of fligh-path angle on cutoff Mach number is shown. Experimental results indicate agreement with the method, since a climb maneuver produced booms of a much decreased intensity as compared with the intensity of those measured in level flight at about the same altitude and Mach number. Comparison of sound pressure levels for the fighter and bomber airp lanes indicated little effect of either airplane size or weight at an altitude of 40,000 feet.

Description: The rather extensive study of the shock losses in transonic compressors can he summarized by the following remarks: 1. A simple flow model can be used to estimate shock losses at the design point for transonic compressor blade rows and results iii reasonable correlation of loss data. It is indicated that shock losses can constitute a sizable portion of the total losses in it transonic compressor rotor. This includes all blade elements at which sonic or higher relative velocities are obtained. 2. Shock losses can he shown to exist across the blade passage (free-stream loss) and by the method of superposition with the blade profile losses result in an estimated design total loss coefficient. 3. The shock configuration was experimentally determined by the rapid pressure rise between the blades as measured by the use of barium titanate crystals. At the minimum loss operating conditions the shock is very similar to that assumed in the simple How model. 4. Shock losses obtained from a more detailed flow model were compared with the losses obtained by the simple flow model. Measured loss distribution from blade to blade closely approaches the analytical shock loss distribution. The measured distribution shows the effect of a shock boundary layer interaction. 5. The analytical method (from the detailed flow model) of determining the shock location ahead of the blade seems to apply reasonably well over a range of incidence angles. The analytical shock losses do not vary a great deal with blade element incidence angles.

Description: Test conditions for the studies are: Mach number varying continuously from approximately 0.8 to 1.1 and Reynolds number (based on maximum diameter of Atlas) approximately 0.451 x 10(exp 6). Camera speed is 2000 frames per second.

Description: The film shows flow over blunt body alone, with internal spike, and with external spikes.

Description: Tests were conducted on several types of porous parachutes, a paraglider, and a simulated retrorocket. Mach numbers ranged from 1.8-3.0, porosity from 20-80 percent, and camera speeds from 1680-3000 feet per second (fps) in trials with porous parachutes. Trials of reefed parachutes were conducted at Mach number 2.0 and reefing of 12-33 percent at camera speeds of 600 fps. A flexible parachute with an inflatable ring in the periphery of the canopy was tested at Reynolds number 750,000 per foot, Mach number 2.85, porosity of 28 percent, and camera speed of 36oo fps. A vortex-ring parachute was tested at Mach number 2.2 and camera speed of 3000 fps. The paraglider, with a sweepback of 45 degrees at an angle of attack of 45 degrees was tested at Mach number 2.65, drag coefficient of 0.200, and lift coefficient of 0.278 at a camera speed of 600 fps. A cold air jet exhausting upstream from the center of a bluff body was used to simulate a retrorocket. The free-stream Mach number was 2.0, free-stream dynamic pressure was 620 lb/sq ft, jet-exit static pressure ratio was 10.9, and camera speed was 600 fps.

Description: Flexible parachute models reefed to one-eighth, one-fourth, one-third, and four tenths of its diameter were towed at speeds of Mach 1.80, 2.00, 2.20 and 2.87. Towline lengths tested were 23.40, 24.38, 26.81, and 29.25 inches. High-speed Schlieren movies of the flow are shown.

Description: An iteration method is presented by which the detailed aerodynamic loading and twist characteristics of a flexible wing with known elastic properties may be calculated. The method is applicable at Mach numbers approaching 1.0 as well as at subsonic Mach numbers. Calculations were made for a wing-body combination; the wing was swept back 45 deg and had an aspect ratio of 4. Comparisons were made with experimental results at Mach numbers from.0.80 to 0.98.

Description: Even though a great deal of theoretical and experimental information has been obtained in recent years on the flow over simple shapes in hypersonic flow a great deal of confusion still exists on how to interpret and extrapolate the results obtained. This paper offers information recently obtained at Langley at Mach numbers ranging from 7 to 21 encompassing both work in air and helium on shapes ranging from rods to delta wings. The results indicate that in most cases methods for making useful estimates of pressure are in hand for simple shapes. However, three-dimensional effects and the interaction between the components considerably complicates the flow fields over delta wings at low angles of attack.

Description: A wind-tunnel investigation has been made on modified-square and circular cylinders to determine the effects of fineness ratio and Reynolds numbers on the crosswind drag characteristics. Fineness ratios from 2 to 14 were investigated over a Reynolds number range from approximately 300,000 to 1,650,000 which corresponded to Mach numbers from 0.057 to 0.377.The result of the investigation show that at supercraft Reynolds numbers the drag coefficient of the circular cylinder increases with increasing Reynolds number for all fineness ratios but at low fineness ratios this effect is considerably less than at higher fineness ratios. For circular cylinders in the high fineness-ratio range there is a reduction in drag as the fineness ratio is decreased except for Reynolds numbers of 900,000 and 1,000,000, whereas at low fineness ratios the opposite trend generally occurs. The addition of hemispherical ends to the circular cylinder gave a substantial decrease in drag at a fineness ratio of 3.27 but the effect was negligible at fineness ratios of 5.27 and 10. The finite-length modified-square cylinder gave the reduction in drag over the two-dimensional modified-square cylinder for the complete range of test Reynolds numbers with the lowest fineness ratio giving the lowest drag at Reynolds numbers above 3O0,OOO.

Description: A concept for interrelating the wave drags of wing-body combinations at supersonic speeds with axial developments of cross-sectional area is presented. A swept-wing-indented-body combination designed on the basis of this concept to have significantly improved maximum lift-drag ratios over a range of transonic and moderate supersonic speeds is described. Experimental results have been obtained for this configuration at Mach numbers from 0.80 to 2.01. Maximum lift-drag ratios of approximately 14 and 9 were measured at Mach numbers of 1.15 and 1.41, respectively.

Description: The inverse method, with the shock wave prescribed to be an elliptic cone at a finite angle of incidence, is applied to calculate numerically the supersonic perfect-gas flow past conical bodies not having axial symmetry. Two formulations of the problem are employed, one using a pair of stream functions and the other involving entropy and components of velocity. A number of solutions are presented, illustrating the numerical methods employed, and showing the effects of moderate variation of the initial parameters.