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Corrugated-core sandwich structures with integrated acoustic resonator arrays have been of recent interest for launch vehicle noise control applications. Previous tests and analyses have demonstrated the ability of this concept to increase sound absorption and reduce sound transmission at low frequencies. However, commercial aircraft manufacturers often require fibrous or foam blanket treatments for broadband noise control and thermal insulation. Consequently, it is of interest to further explore the noise control benefit and trade-offs of structurally integrated resonators when combined with various degrees of blanket noise treatment in an aircraft-representative cylindrical fuselage system. In this study, numerical models were developed to predict the effect of broadband and multi-tone structurally integrated resonator arrays on the interior noise level of cylindrical vibroacoustic systems.

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.

Exposure estimates inside space vehicles, surface habitats, and high altitude aircraft exposed to space radiation are highly influenced by secondary neutron production. The deterministic transport code HZETRN has been identified as a reliable and efficient tool for such studies, but improvements to the underlying transport models and numerical methods are still necessary. In this paper, the forward-backward (FB) and directionally coupled forward-backward (DC) neutron transport models are derived, numerical methods for the FB model are reviewed, and a computationally efficient numerical solution is presented for the DC model. Both models are compared to the Monte Carlo codes HETC-HEDS and FLUKA, and the DC model is shown to agree closely with the Monte Carlo results.

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.

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.

Research committed by the Langley Research Center through 1995 resulting in the HZETRN code provides the current basis for shield design methods according to NASA STD-3000 (2005). With this new prominence, the database, basic numerical procedures, and algorithms are being re-examined with new methods of verification and validation being implemented to capture a well defined algorithm for engineering design processes to be used in this early development phase of the Bush initiative. This process provides the methodology to transform the 1995 HZETRN research code into the 2005 HZETRN engineering code to be available for these early design processes. In this paper, we will review the basic derivations including new corrections to the codes to insure improved numerical stability and provide benchmarks for code verification.

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.

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.

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.

The effects of both the cosmic ray heavy ion exposures and the intense trapped electron exposures are examined with respect to impact on cellular system survival on exterior spacecraft surfaces as well as at interior (shielded) locations for a sample mission to Jupiter’s moons. Radiation transport through shield materials and subsequent exposures are calculated with the established Langley heavy ion and electron deterministic codes. In addition to assessing fractional DNA single and double strand breaks, a variety of cell types are examined that have greatly differing radio-sensitivities. Finally, implications as to shield requirements for controlled biological experiments are discussed.

Special exposure limit recommendations have been designated by the National Council on Radiation Protection (NCRP) for U. S. astronauts in low earth orbit (LEO) operations. These have been established from consideration of a 3% lifetime excess risk of cancer mortality for a 10-yr. active career. The most recent recommendations of the National Council on Radiation Protection and Measurements (NCRP) have incorporated modified procedures for evaluating exposures with accompanying adjustments in career limits. Of special importance are the limit specifications for female exposures, which are approximately 40% less than those for males. Furthermore, radiosensitive organs unique to females require additional attention.

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.

Prior studies have been performed where basic fabric lay-ups of the current Shuttle spacesuit were tested for radiation shielding capabilities. It was found that the fabric portions of the suit give far less protection from radiation than previously estimated. This is due to the porosity and non-uniformity of the fabrics and LCVG components. These findings were incorporated into the spacesuit model developed at NASA Langley Research Center to estimate exposures for mission planning and evaluation of safety during radiation field disturbance. Overall material transmission properties were also less than optimal. In order to evaluate the radiation protection characteristics of some proposed new spacesuit materials, fifteen test target combinations of current baseline and new proposed spacesuit materials were exposed to a low-energy proton beam at Lawrence Berkeley National Laboratory. Each target combination contained all of the necessary spacesuit layers, i.e.

High-speed computational procedures for space radiation shielding have relied on asymptotic expansions in terms of the off-axis scatter and replacement of the general geometry problem by a collection of flat plates. This type of solution was derived for application to human rated systems in which the radius of the shielded volume is large compared to the off-axis diffusion limiting leakage at lateral boundaries. Over the decades these computational codes are relatively complete and lateral diffusion effects are now being added. The analysis for developing a practical full 3D space shielding code is presented.

Traditionally radiation protection is left for evaluation after the completion of other engineering design processes followed by design changes to improve protection leading to off-optimum solutions of design problems. This project is a first attempt to develop optimization procedures with radiation constraint components from the beginning of the design process allowing performance optimization at reduced costs. The traditional limitation of radiation constraint analysis has been the slow computation time and the main focus thus far has been to apply high-performance computing to shielding analysis in preparation for MDO processes. We will describe the problem formulation, the framework for optimization, and progress towards developing highspeed computational procedures.

The great cost of added radiation shielding is a potential limiting factor in many deep space missions. For this enabling technology, we are developing tools for optimized shield design over multi-segmented missions involving multiple work and living areas in the transport and duty phase of various space missions. The total shield mass over all pieces of equipment and habitats is optimized subject to career dose and dose rate constraints. Preliminary studies of deep space missions indicate that for long duration space missions, improved shield materials will be required. The details of this new method and its impact on space missions and other technologies will be discussed. This study will provide a vital tool for evaluating Gateway designs in their usage context. Providing protection against the hazards of space radiation is one of the challenges to the Gateway infrastructure designs.

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.