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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.

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 Mars Reconnaissance Orbiter (MRO) launched on August 12, 2005 and started aerobraking at Mars in March 2006. During the spacecraft's design phase, thermal models of the solar panels and instruments were developed to determine which components would be the most limiting thermally during aerobraking. Having determined the most limiting components, (from a temperature limit standpoint), thermal limits in terms of heat rate were established. Advanced thermal modeling techniques were developed utilizing Thermal Desktop and Patran Thermal. Heat transfer coefficients were calculated using a Direct Simulation Monte Carlo technique. Analysis established that the solar panels were the most limiting components during the aerobraking phase of the mission.

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.

One goal of space radiation research is to reduce the computational time and increase the accuracy of various radiation calculations to aid in their use in a collaborative engineering environment. For example, a fast turn around time is a feature needed for comparison of radiation shielding effects associated with various design configurations of the International Space Station. Research toward this effort has been conducted on various forms of the low energy neutron Boltzmann equation. Simplified models involving the straight ahead approximation, which have fast computational speeds, have been developed at NASA Langley Research Center during the late 1980's as part of a larger high energy ion transport code. Various modifications to improve the accuracy of these computer codes have been an ongoing project. The goal to increase the accuracy of low energy neutron transport without effecting the fast computational times has been a successful ongoing research effort.

The long term exposure of astronauts on the developing International Space Station (ISS) requires an accurate knowledge of the internal exposure environment for human risk assessment and other onboard processes. The natural environment is moderated by the solar wind, which varies over the solar cycle. The HZETRN high charge and energy transport code developed at NASA Langley Research Center can be used to evaluate the neutron environment on ISS. A time dependent model for the ambient environment in low earth orbit is used. This model includes GCR radiation moderated by the Earth’s magnetic field, trapped protons, and a recently completed model of the albedo neutron environment formed through the interaction of galactic cosmic rays with the Earth’s atmosphere. Using this code, the neutron environments for space shuttle missions were calculated and comparisons were made to measurements by the Johnson Space Center with onboard detectors.

The projected radiation levels within the International Space Station (ISS) have been criticized by the Aerospace Safety Advisory Panel in their report to the NASA Administrator. Methods for optimal reconfiguration and augmentation of the ISS shielding are now being developed. The initial steps are to develop reconfigurable and realistic radiation shield models of the ISS modules, develop computational procedures for the highly anisotropic radiation environment, and implement parametric and organizational optimization procedures. The targets of the redesign process are the crew quarters where the astronauts sleep and determining the effects of ISS shadow shielding of an astronaut in a spacesuit. The ISS model as developed will be reconfigurable to follow the ISS. Swapping internal equipment rack assemblies via location mapping tables will be one option for shield optimization.

The hazards of ionizing radiation in space continue to be a limiting factor in the design of spacecraft and habitats. Shielding against such hazards adds to the mission costs and is even an enabling technology in human exploration and development of space. We are developing a web accessible system for radiation hazard evaluation in the design process. The framework for analysis and collaborative engineering is used to integrate mission trajectory, environmental models, craft materials and geometry, system radiation response functions, and mission requirements for evaluation and optimization of shielding distribution and materials. Emphasis of the first version of this integrated design system will address low Earth orbit allowing design system validation using STS, Mir, and ISS measurements. The second version will include Mars, lunar, and other deep space mission analysis.

An experimental investigation was conducted to study the effectiveness of Synthetic Vision Systems (SVS) flight displays as a means of eliminating Low Visibility Loss of Control (LVLOC) and Controlled Flight Into Terrain (CFIT) accidents by low time general aviation (GA) pilots. A series of basic maneuvers were performed by 18 subject pilots during transition from Visual Meteorological Conditions (VMC) to Instrument Meteorological Conditions (IMC), with continued flight into IMC, employing a fixed-based flight simulator. A total of three display concepts were employed for this evaluation. One display concept, referred to as the Attitude Indicator (AI) replicated instrumentation common in today's General Aviation (GA) aircraft. The second display concept, referred to as the Electronic Attitude Indicator (EAI), featured an enlarged attitude indicator that was more representative of a “glass display” that also included advanced flight symbology, such as a velocity vector.

Most of the neutrons seen in the habitable environment of spacecraft in LEO are produced in local materials of the spacecraft structures by the impact of the LEO radiation environment. There are two components of the neutron spectra: one produced near the forward direction and a diffuse isotropic component. The forward component satisfies a Volterra equation and is solved by standard marching procedures. The diffuse component is generally of lower energy and nearly isotropically scattered as they diffuse through the spacecraft structures. Leakage at near boundaries marks the diffusion process and solutions are strongly dependent on forward and backward boundaries with minor contributions from lateral diffusion along spacecraft wall structures. The diffuse neutron equation is solved using multigroup methods with impressed forward and backward boundary conditions.

Trapped protons and electrons in the low earth orbit (LEO) environment of the International Space Station (ISS) encountered during extra-vehicular activity (EVA) may contribute significantly to the cumulative exposure sustained by crew during extended stay missions. A recently developed CAD model of the U. S. Shuttle Space Suit is used to define the shielding properties inherent in the space suit. The model incorporates 28 separate components of the suit, with particular attention given to the helmet and backpack assemblies. Proton and electron energy spectra are taken from the NASA AP8 and AE8 environment models for solar maximum and minimum, and a simulated magnetic storm condition is derived from a 3-sigma projection of the nominal condition. Heavy-ion and electron transport codes developed at NASA-Langley are used in conjunction with the variety of space suit materials, including constituents containing metallic and non-metallic compounds as well as organic polymers.

The General Aviation Element of the Aviation Safety Program's Synthetic Vision Systems (SVS) Project is developing technology to eliminate low visibility induced General Aviation (GA) accidents. SVS displays present computer generated 3-dimensional imagery of the surrounding terrain on the Primary Flight Display (PFD) to greatly enhance pilot's situation awareness (SA), reducing or eliminating Controlled Flight into Terrain, as well as Low-Visibility Loss of Control accidents. SVS-conducted research is facilitating development of display concepts that provide the pilot with an unobstructed view of the outside terrain, regardless of weather conditions and time of day. A critical component of SVS displays is the appropriate presentation of terrain to the pilot. An experimental study is being conducted at NASA Langley Research Center (LaRC) to explore and quantify the relationship between the realism of the terrain presentation and resulting enhancements of pilot SA and performance.

Preliminary braking, steering, and tread wear performance results from testing of 26 x 6.6 and 40 x 14 radial-belted and bias-ply aircraft tires at NASA Langley's Aircraft Landing Dynamics Facility (ALDF) are reviewed. These tire tests are part of a larger, ongoing joint NASA/FAA/Industry Surface Traction And Radial Tire (START) Program involving these two different tire sizes as well as an H46 x 18-20 tire size which has not yet been evaluated. Both dry and wet surface conditions were evaluated on two different test surfaces - nongrooved Portland cement concrete and specially constructed, hexagonal-shaped concrete paver blocks. Use of paver blocks at airport facilities has been limited to ramp and taxiway areas and the industry needs a tire friction evaluation of this paving material prior to additional airport pavement installations.

Following the satellite-level thermal vacuum test for the Cloud-Aerosol Lidar and Infrared Pathfinder Satellite Observation spacecraft, project thermal engineers determined that the radiator used to cool the Integrated Lidar Transmitter subsystem during its operation was oversized. In addition, the thermal team also determined that the operational heaters were undersized, thus creating two related problems. Without the benefit of an additional thermal vacuum test, it was necessary to develop and prove by analysis a laser temperature control scheme using the available resources within the spacecraft along with proper resizing of the radiator. A resizing methodology and new laser temperature control scheme were devised that allowed, with a minimum of 20% heater power margin, the operating laser to maintain temperature at the preferable set point. This control scheme provided a solution to a critical project problem.

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.

Human factors egress testing of the HL-20 Personnel Launch System, a reusable flight vehicle for Space Station crew rotation, was conducted in both the vertical (launch) and horizontal (landing) positions using a full-scale model. Ingress and egress of 10-person crews were investigated with volunteers representing a range of heights. For both the vertical and horizontal positions, interior structural keels had little impact on egress times which were generally less than 30 seconds. Wearing Shuttle partial pressure suits required somewhat more egress time than when ordinary flight suits were worn due to the larger helmet of the Shuttle suit.

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 highly inclined orbit of the International Space Station Alpha exhibits significant radiation exposure contributions from the galactic cosmic rays penetrating the earth's magnetic field. In the absence of an accepted method for estimating the corresponding astronaut risk, we examined the attenuation characteristics using conventional LET dependent quality factors (as one means of representing RBE) and a track-structure repair model fit to cell transformation (and inactivation) data in the C3H10T1/2 mouse cell system obtained by T. C. Yang and coworkers for various ion beams. Although the usual aluminum spacecraft shield is effective in reducing dose equivalent with increasing shield thickness, cell transformation rates are increased for thin aluminum shields providing increased risk rather than protection to large shield thickness.

Judicious shielding strategies incorporated in the initial spacecraft design phase for the purpose of minimizing deleterious effects to onboard systems in intense radiation environments will play a major role in ensuring overall mission success. In this paper, we present parametric shielding analyses for the three Jupiter Icy Moons, Callisto, Ganymede, and Europa, as a function of time in orbit at each moon, orbital inclination, and various thicknesses, for low- and high-Z shielding materials. Trapped electron and proton spectra using the GIRE (Galileo Interim Radiation Electron) environment model were generated and used as source terms to both deterministic and Monte Carlo high energy particle transport codes to compute absorbed dose as a function of thickness for aluminum, polyethylene, and tantalum. Extensive analyses are also presented for graded-Z materials.

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.