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A simple wing leading-edge modification has been developed that delays outer wing panel stall, thus maintaining roll damping to higher angles of attack and delaying the onset of autorotation. The stall angle of attack of the outer wing panel has been shown to be a function of the spanwise length of the leading-edge modification. The margin of spin resistance provided by the modification is being explored through flight tests. Preliminary results have been used to evaluate spin resistance in terms of the difference in angle of attack between outer wing panel stall and the maxiumum attainable angle of attack.

An investigation has been conducted in the Langley 30- by 60-Foot Wind Tunnel to evaluate the performance and stability and control characteristics of a full-scale general aviation airplane equipped with a natural-laminar-flow wing. The study focused on the effects of natural laminar flow and boundary layer transition, and on the effects of several wing leading-edge modifications designed to improve the stall resistance of the configuration. Force and moment data were measured over wide angle-of-attack and sideslip ranges and at Reynolds numbers from 1.4 × 106 to 2.1 × 106 based on the mean aerodynamic chord. Additional measurements were made using hot-film and sublimating-chemical techniques to determine the condition of the wing boundary layer, and wool tufts were used to study the wing stalling characteristics. The investigation showed that large regions of natural laminar flow existed on the wing which would significantly enhance the cruise performance of the configuration.

The Vapor Phase Catalytic Ammonia Removal (VPCAR) system is being developed to recycle water for future NASA Exploration Missions [1,2,3,4,5]. Reduced gravity testing of the VPCAR System has been initiated to identify any potential problems with microgravity operation. Two microgravity testing campaigns have been conducted on NASA's C-9B Reduced Gravity Aircraft. These tests focused on the fluid dynamics of the unit's Wiped-Film Rotating Disk (WFRD) evaporator. The experiments used a simplified system to study the process of forming a thin film on a rotating disk. The configuration simulates the application of feed in the VPCAR's WFRD evaporator. The first round of aircraft testing, which was completed in early 2006, indicated that a problem with microgravity operation of the WFRD existed. It was shown that in reduced gravity the VPCAR wiper did not produce a uniform thin film [6]. The film was thicker near the axis of rotation where centrifugal forces are small.

The results of theory/experiment correlative studies at subsonic and supersonic Mach numbers are presented in this paper. These studies were conducted by using theoretical design tools consisting of the Method of Characteristics, newly-developed integral compressible boundary-layer methods for infinitely swept wings, namely, laminar boundary layer with suction, prediction of neutral instability and transition due to amplification of Tollmien-Schlichting (T.S.) waves and crossflow (C.F.), and a method for predicting separating turbulent boundary-layer characteristics. Results of correlations have indicated that the present integral boundary layer methods are quite successful in predicting transition phenomenon both at transonic and supersonic speeds.

Aviation has experienced one hundred years of dynamic growth and change, resulting in the current air transportation system dominated by commercial airliners in a hub and spoke infrastructure. The first fifty years of aviation was a very chaotic, rapid evolutionary process involving disruptive technologies that required frequent adaptation. The second fifty years produced a stable evolutionary optimization of services based on achieving an objective function of decreased costs. In the third wave of aeronautics over the next fifty years, there is the potential for aviation to transform itself into a more robust, scalable, adaptive, secure, safe, affordable, convenient, efficient, and environmentally fare and friendly system.

The most recent goal of our research program was to identify the optimal features of each of three garments to maintain core temperature and comfort under intensive physical exertion. Four males and 2 females between the ages of 22 and 46 participated in this study. The garments evaluated were the MACS-Delphi, Russian Orlan, and NASA LCVG. Subjects were tested on different days in 2 different environmental chamber temperature/humidity conditions (24°C/H∼28%; 35°C/H∼20%). Each session consisted of stages of treadmill walking/running (250W to 700W at different stages) and rest. In general, the findings showed few consistent differences among the garments. The MACS-Delphi was better able to maintain subjects within a skin and core temperature comfort zone than was evident in the other garments as indicated by a lesser fluctuation in temperatures across physical exertion levels.

A model scaled test apparatus has been designed and assembled to simulate supersonic plume/aircraft structure Interaction for the cruise configuration. Preliminary results have been obtained to demonstrate the severity of the associated acoustic fatigue loads. Two rectangular supersonic nozzles with aspect ratios of 7 and 7.7 ware fabricated with internal convergent-divergent contours designed for Mach numbers of 1.35 and 2.00. A large flat plate was located beneath each nozzle at various nozzle height separations. The plate was instrumented to measure surface dynamic pressure and mean wall temperature. Phase averaged schliern measurements revealed the presence of high intensity acoustic emission from the supersonic plume above the plate and directed upstream. This radiation can be associated with the shock noise generation mechanism. Narrow band spectra of wall dynamic pressure show spectral peaks with amplitude levels as high as 1 PSI.

This paper reviews the status of analytical and empirical propeller noise prediction methods with specific emphasis on those that are suitable for General Aviation propellers. Specifically, the paper reviews the capabilities and limitations of methods that are simple enough for ease of use by industry while providing sufficient accuracy to guide the development of new propeller designs or the modification of existing propeller driven airplanes to satisfy increased certification stringency or cabin comfort objectives.

With the resurgence of the General Aviation industry, the incentive to develop new airplanes for the low-end market has increased. Increased production of small airplanes provides the designers and manufacturers the opportunity to incorporate advanced technologies that are not readily retrofitable to existing designs. Spin resistance is one such technology whose development was concluded by NASA during the 1980’s when the production of small airplanes had slipped into near extinction. This paper reviews the development of spin resistance technology for small airplanes with emphasis on wing design. The definition of what constitutes spin resistance and the resulting amendment of the Federal Aviation Regulations Part 23 to enable certification of spin resistant airplanes are also covered.

An extensive general aviation stall/spin research program is underway at the NASA Langley Research Center. Flight tests have examined the effects of tail design, wing leading edge design, mass distribution, and minor airframe modifications on spin and recovery characteristics. Results and observations on test techniques are presented for the first airplane in the program. Configuration changes produced spins varying from easily recoverable slow, steep spins to unrecoverable, fast flat spins.

Numerical solutions have been generated which simulate flow inside an aircraft engine flying at altitude through an ice crystal cloud. The geometry used for this study is the Honeywell Uncertified Research Engine (HURE) which was recently tested in the NASA Propulsion Systems Laboratory (PSL) in January 2018. The simulations were carried out at predicted operating points with a potential risk of ice accretion. The extent of the simulation is from upstream of the engine inlet to downstream past the strut in the core and bypass. The flow solution is produced using GlennHT, a NASA in-house code. A mixing plane approximation is used upstream and downstream of the fan. The use of the mixing plane allows for steady state solutions in the relative frame. The flow solution is then passed on to LEWICE3D for particle trajectory, impact and breakup prediction. The LEWICE3D code also uses a mixing plane approximation at the boundaries upstream and downstream of the fan.

An automatic trim system for reducing the control forces after an engine failure on a light twin has been investigated on the Langley General Aviation Simulator. The system schedules open-loop trim tab deflections as a function of differential propeller slipstream dynamic pressure and freestream dynamic pressure. The system is described and the airplane-system static and dynamic characteristics are documented. Three NASA research pilots evaluated the effectiveness of the system for takeoff and landing maneuvers. A variety of off-nominal system characteristics were studied. The system was judged to be generally beneficial, providing a 2 to 3 point improvement in pilot rating for the tasks used in the evaluations.

This paper presents a review of our current experimental and theoretical understanding of icing feathers and the role that they play in the formation of ice accretions. It covers the following areas: a short review of past research work related to icing feathers; a discussion of the physical characteristics and terminology used in describing icing feathers; the presence of feathers on ice accretions formed in unswept airfoils, especially at SLD conditions; the role that icing feathers play in the formation of ice accretion shapes on swept wings; the formation of icing feathers from roughness elements; theoretical considerations regarding feather formation, feather interaction to form complex icing structures, the role of film dynamics in the formation of roughness elements and the formation of feathers. Hypotheses related to feather formation and feather growth are discussed.

This paper present a summary of the design studies for the suit port proof of concept. The Suit Port reduces the need for airlocks by docking the suits directly to a rover or habitat bulkhead. The benefits include reductions in cycle time and consumables traditionally used when transferring from a pressurized compartment to EVA and mitigation of planetary surface dust from entering into the cabin. The design focused on the development of an operational proof of concept evaluated against technical feasibility, level of confidence in design, robustness to environment and failure, and the manufacturability. A future paper will discuss the overall proof of concept and provide results from evaluation testing including gas leakage rates upon completion of the testing program.

Computational ice shapes were generated on the boundary layer ingesting engine nacelle of the D8 Double Bubble aircraft. The computations were generated using LEWICE3D, a well-known CFD icing post processor. A 50-bin global drop diameter discretization was used to capture the collection efficiency due to the direct impingement of water onto the engine nacelle. These discrete results were superposed in a weighted fashion to generate six drop size distributions that span the Appendix C and O regimes. Due to the presence of upstream geometries, i.e. the fuselage nose, the trajectories of the water drops are highly complex. Since the ice shapes are significantly correlated with the collection efficiency, the upstream fuselage nose has a significant impact on the ice accretion on the engine nacelle. These complex trajectories are caused by the ballistic nature of the particles and are thus exacerbated as particle size increases.

This paper has presented a technique for the determination of an optimal configuration of fuselage mounted piezoelectric actuators for active structural acoustic control of interior noise in aircraft. The technique has demonstrated much potential in preliminary experiments where actuators were configured to couple into the first principal component of the acoustically coupled fuselage vibration. In this test, average reductions of 6 dB at the error microphones and 4 dB at five auxiliary microphones were observed for a pure tone disturbance at the left forward engine pylon of a business jet. This disturbance was used to simulate an oscillating force due to engine unbalance.

An experiment was conducted in the Icing Research Tunnel (IRT) at NASA Glenn Research Center to study the effect of sweep angle and temperature on the formation of ice accretions on a NACA 0012 swept wing at SLD conditions. From a baseline Appendix-C condition with a MVD of 20m the drop size was changed to 110 and 200m for the SLD cases. Casting data, ice shape tracings, time-sequence and photographic data were obtained. Time-sequence photography was taken during each run to capture in real time the formation of the ice accretion. Measurements of the critical distance were obtained.

Noise is a barrier issue for penetration of civil markets by future tiltrotor aircraft. To address this issue, elements of the NASA Short Haul (Civil Tiltrotor) [SH(CT)] program are working in three different areas: noise abatement, noise reduction, and noise prediction. Noise abatement refers to modification of flight procedures to achieve quieter approaches. Noise reduction refers to innovative new rotor designs that would reduce the noise produced by a tiltrotor. Noise prediction activities are developing the tools to guide the design of future quiet tiltrotors. This paper presents an overview of SH(CT) activities in all three areas, including sample results.

Aerodynamic effects induced from vectoring an exhaust jet are investigated using a well established thin-layer Reynolds averaged Navier-Stokes code. This multiple block code has been modified to allow for the specification of jet properties at a block face. The applicability of the resulting code for thrust vectoring applications is verified by comparing numerically and experimentally determined pressure coefficient distributions for a jet-wing afterbody configuration with a thrust-vectoring 2-D nozzle. Induced effects on the body and nearby wing from thrust vectoring are graphically illustrated.

Presented in this paper are the results of three nonlinear computer programs, KRASH, ACTION and DYCAST used to analyze the dynamic response of a twin-engine, low-wing airplane section subjected to a 8.38 m/s (27.5 ft/s) vertical impact velocity crash condition. This impact condition simulates the vertical sink rate in a shallow aircraft landing or takeoff accident. The three distinct analysis techniques for nonlinear dynamic response of aircraft structures are briefly examined and compared versus each other and the experimental data. The report contains brief descriptions of the three computer programs, the respective aircraft section mathematical models, pertinent data from the experimental test performed at NASA Langley, and a comparison of the analyses versus test results. Cost and accuracy comparisons between the three analyses are made to illustrate the possible uses of the different nonlinear programs and their future potential.