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On January 14, 2004 President George W Bush outlined a new vision for NASA that has humans venturing back to the moon by 2020. With this ambitious goal, new tools and models have been developed to help define and predict the amount of space radiation astronauts will be exposed to during transit and habitation on the moon. A representative scenario is used that includes a trajectory from LEO to a Lunar Base, and simplified CAD models for the transit and habitat structures. For this study galactic cosmic rays, solar proton events, and trapped electron and proton environments are simulated using new dynamic environment models to generate energetic electron, and light and heavy ion fluences. Detailed calculations are presented to assess the human exposure for transit segments and surface stays.

Pressure Sensitive Paint (PSP) has been used for several years by the aircraft industry in transonic wind tunnel testing where the oxygen concentrations are low and the luminescence of the paint is easily recorded. Extending PSP to slower speeds where the oxygen concentrations are closer to atmospheric conditions is much more challenging. For the past few years, work has been underway at both Wright Patterson Air Force Base and Ford Motor Company to advance PSP techniques for testing at slower speeds. The CRADA (Cooperative Research and Development Agreement) provided a way for comparisons to be made of the different PSP systems that were being investigated. This paper will report on PSP tests conducted as part of the CRADA.

A discussion is given of the ongoing research related to laminar flow airfoils, nacelles, and wings where the laminar flow is maintained by a favorable pressure gradient, surface suction or a combination of the two. Design methologies for natural laminar flow airfoil sections and wings for both low and high speed applications are outlined. Tests of a 7-foot chord, 23° sweep laminar-flow-control-airfoil at high subsonic Mach numbers are described along with the associated stability theory used to design the suction system. The state-of-the-art of stability theory is simply stated and a typical calculation illustrated. In addition recent computer simulations of transition using the time dependent Navier-Stokes (N-S) equations are briefly described. Advances in wind tunnel capabilities and instrumentation will be reviewed followed by the presentation of a few results from both wind tunnels and flight. Finally, some suggestions for future work will complete the paper.

Solar proton events (SPEs) represent the single-most significant source of acute radiation exposure during space missions. Historically, an exponential in rigidity (particle momentum) fit has been used to express the SPE energy spectrum using GOES data up to 100 MeV. More recently, researchers have found that a Weibull fit better represents the energy spectrum up to 1000 MeV (1 GeV). In addition, the availability of SPE data extending up to several GeV has been incorporated in analyses to obtain a more complete and accurate energy spectrum representation. In this paper we discuss the major SPEs that have occurred over the past five solar cycles (~50+ years) in detail - in particular, Aug 1972 and Sept & Oct 1989 SPEs. Using a high-energy particle transport/dose code, radiation exposure estimates are presented for various thicknesses of aluminum. The effects on humans and spacecraft systems are also discussed in detail.

A scramjet exhaust simulation technique for hypersonic wind tunnel testing has been developed. Mixtures of Argon and Freon correctly match the inviscid simulation parameters of Mach number, static-pressure ratio, and the ratio of specific heats at the combustor exit location; this simulation is accomplished at significantly reduced temperatures and without combustion. An investigation of nozzle parametrics in a Mach 6 freestream showed that the external nozzle ramp angle, the cowl trailing-edge angle, an external nozzle flow fence and the nozzle static-pressure ratio significantly affected the external nozzle thrust and pitching moment as measured by the integration of surface-pressure data. A comparison of Argon-Freon and air exhaust simulation showed that the external nozzle thrust and pitching moment were in error by roughly a factor of 2 using air due to the incorrect match of the ratio of specific heats.

A new method is developed for calculating the low energy neutron flux in a space environment which is protected from galactic cosmic rays (GCR) and solar particle events (SPE) by shielding materials. Our calculations are compared with low energy neutron flux flight data recorded on four different STS low earth orbit missions. We also compare our neutron flux calculations with the low energy neutron flux data recorded by MIR. The low energy neutron flux calculations can be described as a deterministic method for solving the Boltzmann equation for the light ion flux associated with a given environment. Existing Monte Carlo neutron flux simulations associated with the MIR and ISS space stations are also compared with our deterministic method for calculating neutron flux.

This paper presents a simplified orbit analysis program developed to calculate orbital parameters for the thermal analysis of spacecraft and space-flight instruments. The program calculates orbit data for inclined and sunsynchronous earth orbits. Traditional orbit analyses require extensive knowledge of orbital mechanics to produce a simplified set of data for thermal engineers. This program was created to perform orbital analyses with minimal input and provides the necessary output for thermal analysis codes. Engineers will find the program to be a valuable analysis tool for fast and simple orbit calculations. A description of the program inputs and outputs is included. An overview of orbital mechanics for inclined and Sun-synchronous orbits is also presented. Finally, several sample cases are presented to illustrate the thermal analysis applications of the program.

Reynolds number effects noted from selected test programs conducted in the Langiey 0.3-Meter Transonic Cryogenic Tunnel (0.3-m TCT) are discussed. The tests, which cover a unit Reynolds number range from about 2.0 to 80.0 million per foot, summarize effects of Reynolds number on: 1) aerodynamic data from a supercritical airfoil, 2) results from several wall interference correction techniques, and 3) results obtained from advanced, cryogenic test techniques. The test techniques include 1) use of a cryogenic sidewall boundary layer removal system, 2) detailed pressure and hot wire measurements to determine test section flow quality, and 3) use of a new hot film system suitable for transition detection in a cryogenic wind tunnel. The results indicate that Reynolds number effects appear most significant when boundary layer transition effects are present and at high lift conditions when boundary layer separation exists on both the model and the tunnel sidewall.

In view of manned missions targeted to the Moon, 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 determination of the particle flux and spectra at any time and at any point of the lunar surface. With this goal in mind, a new model of the Moon’s radiation environment due to Galactic Cosmic Rays (GCR) and Solar Particle Events (SPE) has been developed. Primary particles reach the lunar surface, and are transported all throughout the subsurface layers, with backscattering patterns taken into account. The surface itself has been modeled as regolith and bedrock, with composition taken from the results of the instruments flown on the Apollo missions, namely on the Apollo 12 from the Oceanus Procellarum landing site. Subsurface environments like lava tubes have been considered in the analysis.

An advanced analytical methodology has been developed for predicting the residual strength of stiffened thin-sheet riveted shell structures such as those used for the fuselage of a commercial transport aircraft. The crack-tip opening angle elastic-plastic fracture criterion has been coupled to a geometric and material nonlinear finite element shell code for analyzing complex structural behavior. An automated adaptive mesh refinement capability together with global-local analysis methods have been developed to predict the behavior of fuselage structure with long cracks. This methodology is currently being experimentally verified. Advanced nondestructive inspection technology has been developed that will provide airline operators with the capability to conduct reliable and economical broad-area inspections of aircraft structures.

An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5% scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from α = -5 to 85 deg. and β = -45 to 45 deg. at a Reynolds number of 0.24x10⁶ and Mach number of 0.06. The 3.5% scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5% scale GTM.

A summary of information on airplane aerodynamics, ground effects, propulsion system aerodynamics, stability and control, and flying qualities of jet V/STOL airplanes - both direct jet lift and lift fan configurations is presented. The information is applicable to high-speed fighter-type airplanes. Research work in the following areas is reviewed: 1. Wind tunnel and other experimental research on jet-induced effects (including ground effects) on the aerodynamics and stability and control in the VTOL, STOL, hovering, and transition ranges of flight. 2. Experimental research on propulsion aerodynamics in the hovering and very low speed ranges of flight. 3. Flight-test experience on the flying qualities of several jet V/STOL airplanes.

The Aircraft Landing Dynamics Facility (ALDF), formerly called the Landing Loads Track, is described. The paper gives a historical overview of the original NASA Langley Research Center Landing Loads Track and discusses the unique features of this national test facility. Comparisions are made between the original track characteristics and the new capabilities of the Aircraft Landing Dynamics Facility following the recently completed facility update. Details of the new propulsion and arresting gear systems are presented along with the novel features of the new high-speed carriage. The data acquisition system is described and the paper concludes with a review of future test programs.

Fuel costs comprise a major portion of air transport operating costs. Thus, energy efficiency is an essential design goal for future transport aircraft. Advanced composite structures, advanced wing geometries, and active control systems all promise substantial benefits in fuel efficiency and direct operating cost for derivative and new aircraft introduced by 1985. Technology for maintenance of a laminar boundary layer in cruise offers great benefits in fuel efficiency and direct operating cost and may be ready for application to transports introduced in the 1990's. NASA and the air transport industry are cooperating in a comprehensive Aircraft Energy Efficiency Program to expedite the introduction of these advanced technologies into production aircraft.

Alternate Environmental Control and Life Support System (ECLSS) technologies were evaluated to reduce Space Station resources and dependence on expendables resupplied from Earth to sustain a multiperson crew in low-Earth orbit. Options were evaluated to close the oxygen (O2) loop by removing carbon dioxide (CO2) from the cabin air, reducing the CO2 to water, and electrolyzing the water to provide metabolic O2 for crew consumption. Options were also evaluated to close the urine/flush, condensate, and hygiene water loops to provide potable water for crew use. Specific evaluation parameters were derived which included weight, power, volume, maintenance, resupply consumables, and technology readiness.

The effect of different types of control force actuator models and geometries on the intensity flow between a cylindrical shell and the contained acoustic space has been analytically investigated. The primary source was an external monopole located adjacent to the exterior surface of the cylinder midpoint. Actuator models of normal point forces and in-plane piezoelectric patches were assumed attached to the wall of a simply-supported, elastic cylinder closed with rigid end caps. Control inputs to the actuators were determined such that the integrated square of the pressure over the interior of the vibrating cylinder was a minimum. Test cases involving a resonant acoustic response and a resonant structural response were investigated. Significant interior noise reductions were achieved for all actuator configurations. Intensity distributions for the test cases show the circuitous path of structural acoustic power flow.

An experimental investigation of the flow over the rear end of a 0.16 scale notchback automobile configuration has been conducted in the NASA Langley Basic Aerodynamics Research Tunnel (BART). The objective of this work was to investigate the flow separation that occurs behind the backlight and obtain experimental data that can be used to understand the physics and time-averaged structure of the flow field. A three-component laser velocimeter was used to make non-intrusive, velocity measurements in the center plane and in a single cross-flow plane over the decklid. In addition to off-body measurements, flow conditions on the car surface were documented via surface flow visualization, boundary layer measurements, and surface pressures.

A research program is in progress to study the effects of the propeller slipstream on natural laminar flow. Flight and wind tunnel measurements of the wing boundary layer have been made using hot-film velocity sensor probes. The results show the boundary layer, at any given point, to alternate between laminar and turbulent states. This cyclic behavior is due to periodic external flow turbulence originating from the viscous wake of the propeller blades. Analytic studies show the cyclic laminar/turbulent boundary layer layer to result in a significantly lower wing section drag than a fully turbulent boundary layer. The application of natural laminar flow design philosophy yields drag reduction benefits in the slipstream affected regions of the airframe, as well as the unaffected regions.

Recent studies have found that high mass concentrations of ice particles in regions of deep convective storms can adversely impact aircraft engine and air probe (e.g. pitot tube and air temperature) performance. Radar reflectivity in these regions suggests that they are safe for aircraft penetration, yet high ice water content (HIWC) is still encountered. The aviation weather community seeks additional remote sensing methods for delineating where ice particle (or crystal) icing conditions are likely to occur, including products derived from geostationary (GEO) satellite imagery that is now available in near-real time at increasingly high spatio-temporal detail from the global GEO satellite constellation.

Recently, the concept of the application of hybrid laminar flow to modern commercial transport aircraft was successfully flight tested on a Boeing 757 aircraft. In this limited demonstration, in which only part of the upper surface of the swept wing was designed for the attainment of laminar flow, significant local drag reduction was measured. This paper addresses the potential application of this technology to laminarize the external surface of large, modern turbofan engine nacelles which may comprise as much as 5-10 percent of the total wetted area of future commercial transports. A hybrid-laminar-fiow-control (HLFC) pressure distribution is specified and the corresponding nacelle geometry is computed utilizing a predictor/corrector design method. Linear stability calculations are conducted to provide predictions of the extent of the laminar boundary layer. Performance studies are presented to determine potential benefits in terms of reduced fuel consumption.