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A flight test investigation has been conducted to determine the performance of wingtip vortex turbines and their effect on aircraft performance. The turbines were designed to recover part of the large energy loss (induced drag) caused by the wingtip vortex. The turbine, driven by the vortex flow, reduces the strength of the vortex, resulting in an associated induced drag reduction. A four-blade turbine was mounted on each wingtip of a single-engine, T-tail, general aviation airplane. Two sets of turbine blades were tested, one with a 15° twist (washin) and one with no twist. The power recovered by the turbine and the installed drag increment were measured. A trade-off between turbine power and induced drag reduction was found to be a function of turbine blade incidence angle. This test has demonstrated that the wingtip vortex turbine is an attractive alternate, as well as an emergency, power source.

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.

To meet the challenge of future deep space programs an accurate and efficient engineering code for analyzing the shielding requirements against high-energy galactic heavy radiations is needed. Such engineering design codes require establishing validation processes using laboratory ion beams and space flight measurements in realistic geometries. In consequence, a new version of the HZETRN code capable of simulating HZE ions with either laboratory or space boundary conditions is currently under development. The new code, GRNTRN, is based on a Green's function approach to the solution of Boltzmann's transport equation and like its predecessor is deterministic in nature. Code validation in the laboratory environment is addressed by showing that GRNTRN accurately predicts energy loss spectra as measured by solid-state detectors in ion beam experiments.

The Mars Reconnaissance Orbiter (MRO) launched on August 12, 2005 and began aerobraking at Mars in March 2006. In order to save propellant, MRO used aerobraking to modify the initial orbit at Mars. The spacecraft passed through the atmosphere briefly on each orbit; during each pass the spacecraft was slowed by atmospheric drag, thus lowering the orbit apoapsis. The largest area on the spacecraft, most affected by aeroheating, was the solar arrays. A thermal analysis of the solar arrays was conducted at NASA Langley Research Center to simulate their performance throughout the entire roughly 6-month period of aerobraking. A companion paper describes the development of this thermal model. This model has been correlated against many sets of flight data. Several maneuvers were performed during the cruise to Mars, such as thruster calibrations, which involve large abrupt changes in the spacecraft orientation relative to the sun.

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.

Current viscous drag reduction research explores the limits of practical applications of natural laminar flow (NLF) for airplane drag reduction. To better understand these limits, advanced measurement techniques are required to study the characteristics of laminar to turbulent boundary-layer transition. Recent NASA research indicates that the transition mode which involves laminar separation can be detected using arrayed hot-film laminar separation sensor concepts. These surface-mounted sensors can provide information on the location of the laminar separation bubble as well as bubble length. This paper presents two different laminar separation sensor configurations developed in the NASA program and presents results of wind-tunnel and flight evaluations of the sensors as tools to detect boundary-layer transition.

Technology integration studies were made to examine the impact of emerging technologies on fighter aircraft. The technologies examined included advances in aerodynamics, controls, structures, propulsion, and systems and were those which appeared capable of being ready for application by the turn of the century. A primary impetus behind large increases in fighter capability will be the rapid increase in fighter engine thrust-to-weight ratio. High thrust-weight engines, integrated with other advanced and emerging technologies, can result in small extremely maneuverable fighter aircraft that have thrust-weight ratios of 1.4+ and weight one-half as much as today's fighters. Future fighter aircraft requirements are likely to include a turn capability in excess of 7g's throughout much of the maneuver envelope, post-stall maneuverability, STOVL or VTOL, and a single engine for low cost.

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.

NASA and the FAA conducted two flight campaigns to quantify onboard weather radar measurements with in-situ measurements of high concentrations of ice crystals found in deep convective storms. The ultimate goal of this research was to improve the understanding of high ice water content (HIWC) and develop onboard weather radar processing techniques to detect regions of HIWC ahead of an aircraft to enable tactical avoidance of the potentially hazardous conditions. Both HIWC RADAR campaigns utilized the NASA DC-8 Airborne Science Laboratory equipped with a Honeywell RDR-4000 weather radar and in-situ microphysical instruments to characterize the ice crystal clouds. The purpose of this paper is to summarize how these campaigns were conducted and highlight key results. The first campaign was conducted in August 2015 with a base of operations in Ft. Lauderdale, Florida.

This paper reports the status of the NASA Langley Research Center program aimed at the development of the technology required for large-scale Magnetic Suspension and Balance Systems. The use of magnetic suspension of the model in a wind tunnel is seen to be the only viable method to eliminate aerodynamic interference problems arising with mechanical model-supports. The two small-scale magnetic suspension systems in operation at Langley are the only ones now active in the U.S. The general features and capabilities of these two systems and all of the ongoing research in the use of magnetic suspension are described.

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.

Meeting radiation protection requirements during EVA is predominantly an operational issue with some potential considerations for temporary shelter. The issue of spacesuit shielding is mainly guided by the potential of accidental exposure when operational and temporary shelter considerations fail to maintain exposures within operational limits. In this case, very high exposure levels are possible which could result in observable health effects and even be life threatening. Under these assumptions, potential spacesuit radiation exposures have been studied using known historical solar particle events to gain insight on the usefulness of modification of spacesuit design in which the control of skin exposure is a critical design issue and reduction of blood forming organ exposure is desirable.

The development of “next generation” human-rated space vehicles, surface habitats and rovers, and spacesuits will require the integration of low-cost, lightweight materials that also include excellent mechanical, structural, and thermal properties. In addition, it is highly desirable that these materials exhibit excellent space radiation exposure mitigation properties for protection of both the crew and onboard sensitive electronics systems. In this paper, we present trapped electron and proton space radiation exposure computational results for a variety of materials and shielding thicknesses for several earth orbit scenarios that include 1) low earth orbit (LEO), 2) medium earth orbit (MEO), and 3) geostationary orbit (GEO). We also present space radiation exposure (galactic cosmic radiation and solar particle event) results as a function of selected materials and thicknesses.

As part of NASA’s Aviation Safety and Security Program, a simulation study of a twin-jet transport aircraft crew training simulation was conducted to address fidelity for upset or loss-of-control flight conditions. Piloted simulation studies were conducted to compare the baseline crew training simulation model with an enhanced aerodynamic model that was developed for high-angle-of-attack conditions. These studies were conducted in a flaps-up configuration and covered the approach-to-stall, stall and post-stall flight regimes. Qualitative pilot comments and preliminary comparison with flight test data indicate that the enhanced model is a significant improvement over the baseline. Some of the significant unrepresentative characteristics that are predicted by the baseline crew training simulation for flight in the post-stall regime have been identified.

As shielding materials are developed for protection against the hazards of galactic cosmic rays, it is desirable to develop a protocol for rapid assessment of shielding properties. Solid state energy loss detectors are often used to estimate the charge and energy of particles in ion beam experiments. The direct measurement is energy deposited in the detector. As a means of separating the charge components in typical shield transmission studies with observation, a stack of many such detectors is used. With high-energy beams and thin targets, surviving primaries and fragments emerging from the target have nearly-equal velocities and deposited energy scales with the square of the charge, simplifying the data analysis. The development of a transport model for the shield and detector arrangement and evaluation of prediction of the energy loss spectrum for direct comparison with the experimentally derived data allows a rapid assessment of the shield transmission characteristics.

The various sources of loads of importance in the design of launch vehicles are discussed generally. The natural vibration modes, which are an essential ingredient in all loads problems, are described for the Saturn vehicle as obtained from a structural replica model. The effects of fuel loading on the structural modes are considered and the results obtained on the model are compared with full-scale results. Launch-vehicle buffeting is discussed and the buffeting characteristics of a manned lunar vehicle configuration are described. Some evaluations of buffeting scaling laws obtained with this configuration are presented. Estimates of the acoustic environment on the spacecraft associated with engine noise at lift-off and with aerodynamic noise due to buffeting are shown. The steady and vibratory loads due to ground winds as determined from wind-tunnel tests of a Saturn model are presented.

In view of manned missions targeted to Mars, for which radiation exposure is one of the greatest challenges to be tackled, it is of fundamental importance to have available a tool, which allows the determination of the particle flux and spectra at any time at any point of the Martian surface. With this goal in mind, a new model for the radiation environment to be found on the planet Mars due to Galactic Cosmic Rays (GCR) and Solar Particle Events (SPE) has been developed. Primary particle environments computed for Martian conditions are transported within the Mars atmosphere, with temporal properties modeled with variable timescales, down to the surface, with topography and backscattering patterns taken into account. The atmospheric chemical and isotopic composition has been modeled over results from the in-situ Viking Lander measurements for both major and minor components.

High Ice Water Content (HIWC) has been identified as a primary causal factor in numerous engine events over the past two decades. Previous attempts to develop a remote detection process utilizing modern commercial radars have failed to produce reliable results. This paper discusses the reasons for previous failures and describes a new technique that has shown very encouraging accuracy and range performance without the need for any modifications to industry’s current radar design(s). The performance of this new process was evaluated during the joint NASA/FAA HIWC RADAR II Flight Campaign in August of 2018. Results from that evaluation are discussed, along with the potential for commercial application, and development of minimum operational performance standards for future radar products.

This study addresses the possibility of beaming laser power from synchronous lunar orbits (the L1 and L2 LaGrange points) to a manned long-range lunar rover. The rover and two versions of a satellite system (one powered by a nuclear reactor; the other by photovoltaics) are described in terms of their masses, geometry, power needs, mission and technological capabilities. Laser beam power is generated by a laser diode array in the satellite and converted to 30 kW of electrical power at the rover. Present technological capabilities, with some extrapolation to near future capabilities, are used in the descriptions. The advantages of the two satellite/rover systems over other such systems and over rovers with on-board power are discussed along with the possibility of enabling other missions.

Recently, there has been a strong emphasis on aerodynamic and aeroacoustic wind tunnel testing of automobiles. While significant level resources have been spent on investigating aerodynamics, the methodology has not changed appreciably since the beginning of aerodynamics as a science. Over the past decade, a number of global flow diagnostic techniques have been developed that drastically increase the quality and quantity of data from wind tunnel testing. One of these technologies is the use of pressure sensitive luminescent coatings, known as pressure-sensitive paint, a method which has matured considerably since its inception and is now used extensively in aerospace applications with good results. The goal of this research is to implement this technology in the full scale testing of high performance automotive vehicles. This paper discusses the details of a preliminary test, such as technique, paint formulation, camera and lighting hardware, and data reduction and analysis.