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Physiological, psychological, and bacteriological aspects of 20 days in full pressure suits in 100 percent oxygen environment
The development of high-power high-voltage space systems will require advances in power generation and processing. The systems must be reliable, adaptable, and durable for space mission success. The issues, which must be resolved in order to produce a high power system, are weight and volume reduction of components and modules and the creation of a reliable high repetition pulse power processor. Capacitor energy density must be increased by twice the present capacity and packaging must be reduced by a factor of 10 to 20 times. The packaging must also protect the system from interaction with the natural space environment and the induced environment, produced from spacecraft systems and environment interaction.
A major objective of the Space Act of 1958 which led to the establishment of the National Aeronautics and Space Administration (NASA) was the dissemination of science and technology. Today, under NASA administrator Daniel Goldin and the White House, there is a reemphasis on the dissemination and transfer of NASA science and technology to U.S. industry: both aerospace and non aerospace. The goal of this transfer of science and technology is to aid U.S. industries, making them more competitive in the global economy. After 69 months in space, LDEF provided new and important information on the space environment and how this hostile environment impacts spacecraft materials and systems. The space environment investigated by the LDEF researchers included: ionizing radiation, ultraviolet radiation, meteoroid and debris, atomic oxygen, thermal cycling, vacuum, microgravity, induced contamination and various synergistic effects. The materials used as part of LDEF and its experiments include polymers, metals, glass, paints and coatings. Fiber optic, mechanical, electrical, and optical systems were also used on LDEF. As part of the effort to disseminate and transfer LDEF science and technology, an annotated bibliographic database is being developed. This bibliography will be available electronically, as well as in hard copy. All LDEF domestic and foreign publications in the open literature, including scientific journals, the NASA LDEF Symposia volumes, books, technical reports and unrestricted contractor reports will be included in this database. The hard copy, as well as the electronic database, will be categorized by section in the scientific and technical discipline.
As the commercial spacecraft industry increases in size, there will be a corresponding increase in the number of manned spacecraft built and operationally flown each year. Industry teams for these new spacecraft will have multiple design and operational choices to make for each of these spacecraft s subsystems. The carbon dioxide (CO2) removal subsystem of the environmental control and life support system is one that presents such challenges. This paper seeks to aid industry in making design and operations choices by providing a document containing lessons learned by the Space Shuttle Program s Operations team, with specific focus given to the non-regenerable CO2 removal system currently used by the Space Shuttle. Carbon dioxide, one of the key byproducts of respiration, can lead to injury and death if allowed to build up in a spacecraft s habitable environment. Therefore, any spacecraft s environmental control and life support system must contain a method for removing this hazard. These removal systems can either be non-regenerable or regenerable. While this paper defines the difference between these types and presents a generic comparison of their capabilities, the focus is specifically on the Space Shuttle s CO2 removal systems. This will include a short discussion of the Space Shuttle s regenerable amine solid absorption system that was part of the Extended Duration Orbiter (EDO) modification, however, emphasis is given to the non-regenerable Lithium Hydroxide canister system used as the prime removal method employed by the Space Shuttle, including a discussion on why this method was chosen over a regenerable system. A full exposition is given on all of the considerations required for mission planning when using a non-regenerable CO2 removal system. Key discussion items include: airflow lessons learned, recent physiological issues related to short term and long term exposure and how lower levels may be more harmful than previously thought, lithium hydroxide dust issues, mass vs. stowage comparisons, cost, complexity considerations, operational margin considerations, timing considerations with respect to the crew s schedule, and logistical management. The conclusion of the paper finalizes the discussion and contains a listing of considerations in regards to CO2 management using a non-regenerable removal system in the near future.
Satellites in geosynchronous orbits have been found to be charged to significant negative voltages during encounters with geomagnetic substorms. When satellite surfaces are charged, there is a probability of enhanced contamination from charged particles attracted back to the satellite by electrostatic forces. This could be particularly disturbing to large satellites using sensitive optical systems. In this study the NASA Charging Analyzer Program (NASCAP) is used to evaluate qualitatively the possibility of such enhanced contamination on a conceptual version of a large satellite. The evaluation is made by computing surface voltages on the satellite due to encounters with substorm environments and then computing charged-particle trajectories in the electric fields around the satellite. Particular attention is paid to the possibility of contaminants reaching a mirror surface inside a dielectric tube because this mirror represents a shielded optical surface in the satellite model used. Deposition of low energy charged particles from other parts of the spacecraft onto the mirror was found to be possible in the assumed moderate substorm environment condition. In the assumed severe substorm environment condition, however, voltage build up on the inside and edges of the dielectric tube in which the mirror is located prevents contaminants from reaching the mirror surface.
The interaction of the aerospace environment with spacecraft surfaces and onboard, high voltage spacecraft systems operating over a wide range of altitudes from low Earth orbit to geosynchronous orbit is considered. Emphasis is placed on control of spacecraft electric potential. Electron and ion beams, plasma neutralizers material selection, and magnetic shielding are among the topics discussed.
Thermal radiation characteristics of spacecraft temperature control louvers in solar space environment, discussing specular reflection
People and spacecraft exploring the Moon face unforgiving environments. The lunar surface is covered in fine, abrasive dust churned by meteoroid impacts. Explorers face extreme thermal, radiation, and lighting conditions. Landing spacecraft must account for erosion caused by rocket exhaust, which can be hazardous for the vehicle or nearby assets. This talk will survey these challenges and the work being done to overcome them to achieve mission success.
The neutron environment inside spacecraft has been of interest from a scientific and radiation protection perspective since early in the history of manned spaceflight. With 1:.1e exception of a few missions which carried plutonium-fueled radioisotope thermoelectric generators, all of the neutrons inside the spacecraft are secondary radiations resulting from interactions of high-energy charged particles with nuclei in the Earth's atmosphere, spacecraft structural materials, and the astronaut's own bodies. Although of great interest, definitive measurements of the spacecraft neutron field have been difficult due to the wide particle energy range and the limited available volume and power for traditional techniques involving Bonner spheres. A multitude of measurements, however, have been made of the neutron environment inside spacecraft. The majority of measurements were made using passive techniques including metal activation fo ils, fission foils, nuclear photoemulsions, plastic track detectors, and thermoluminescent detectors. Active measurements have utilized proton recoil spectrometers (stilbene), Bonner Spheres eRe proportional counter based), and LiI(Eu)phoswich scintillation detectors. For the International Space Station (ISS), only the plastic track! thermoluminescent detectors are used with any regularity. A monitoring program utilizing a set of active Bonner spheres was carried out in the ISS Lab module from March - December 200l. These measurements provide a very limited look at the crew neutron exposure, both in time coverage and neutron energy coverage. A review of the currently published data from past flights will be made and compared with the more recent results from the ISS. Future measurement efforts using currently available techniques and those in development will be also discussed.
The issue of the effects of the space environment on spacecraft needs to be understood for the long term exposure of structures in space. In order to better understand the effect of these hostile phenomena on spacecraft, several types of studies are worth performing in order to simulate at some level the effect of the environment. For example the effect of protons and electrons impacting structural materials are easily simulated through experiments using the Van de Graff and Pelletron accelerators currently housed at MSFC. Proton fluxes with energies of 700 KeV - 2.5 MeV can be generated and used to impinge on sample targets to determine the effects of the particles. Also the Environmental Effects Facility at MSFC has the capability to generate electron beams with energies from 700 KeV to 2.5 MeV. These facilities will be used in this research to simulate space environmental effects from energetic particles. Ultraviolet radiation, particularly less than 400 nm wavelength, is less well characterized at this time. The Environmental Effects Facility has a vacuum system dedicated to studying the effects of ultraviolet radiation on specific surface materials. This particular system was assembled in a previous study in order to perform a variety of experiments on materials proposed for the Space Station. That system has continued to function as planned and has been used in carrying out portions of the proposed study.
The Long Duration Exposure Facility (LDEF) was retrieved from Earth orbit in January 1990 after spending almost six years in space. It had flown in a near-circular orbit with an inclination of 28.5 degrees. Initially, the orbit altitude was approximately 257 nautical miles; however, when the LDEF was retrieved the orbit altitude had decayed to approximately 179 nautical miles. The LDEF was passively stabilized about three axes while in free flight, making it an ideal platform for exposing experiments which were measuring the environments of near-Earth space and investigating the long-term effects of these environments on spacecraft. A brief overview of the encountered environments that were of most interest to the LDEF investigators is presented.
Abstract: Spacecraft thermal design presents unique thermal and fluid challenges caused by a variety of harsh spacecraft thermal environments. In addition, at cryogenic temperatures (i.e., less than 123 K (-150° C), and sometimes as low as 50 milliKelvin), more challenges arise, such as the increasing importance of mitigating thermal radiation, the effects of coefficient of thermal expansion mismatches between different materials, changes in the physics behavior of materials, and increasing thermodynamic difficulty of lifting heat up large temperature gradients. In this talk, Dr. Rueger will present a broad overview of spacecraft thermal environments; cryogenic design practices; components of interest for the Cryogenics and Fluids Branch (e.g., cryocoolers, adiabatic demagnetization refrigerators, heat switches, and superconducting leads); cryogenic considerations in some current NASA mission examples (e.g., the DragonFly mission to Titan and the X-ray Imaging and Spectroscopy Mission); and capabilities of the Cryogenics and Fluids Branch at NASA.
The development of spacecraft with electrical loads that require high voltage power is discussed. The high voltage solar array has been considered for supplying d.c. power directly to high voltage loads such as ion thrusters and communication tubes without intermediate power processing. Space power stations for transferring solar power to earth are being studied in the 40 kilovolt, multikilowatt regime. Analytical and experimental studies have determined that with the advent of high voltage power, new problems will arise through the interaction of the high voltage surfaces with the charged particle environment of space. The interactive environment has been identified and duplicated to some extent in simulation facilities at NASA-Lewis Research Center and at several contractor locations.
Thermal radiation characteristics of spacecraft temperature control louvers in solar space environment, discussing specular reflection
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Estimates of short and long term solar activity are required for calculating variations in the environment with regard to spacecraft charging, radiation effects, and orbital lifetime. Correlations appear to exist between the time of solar activity and the time of operational anomalies due to electrical discharges when the dielectric surfaces of geosynchronous spacecraft are charged by interaction with the ambient plasma to levels above breakdown voltage. An ability to predict the solar induced variation variability of the plasmas could permit refinement of design criteria. The influence of the radiation on spacecraft materials, systems, and manned operations is summarized. Solar radition effects on the orbital altitude atmospheric density environment and spacecraft lifetimes are also considered.
Analytical techniques and calculations in planetary quarantine and spacecraft sterilization
Predicting the rate of occurrence of single event effects (SEEs) in space requires knowledge of the radiation environment and the response of electronic devices to that environment. Several analytical models have been developed over the past 36 years to predict SEE rates. The first error rate calculations were performed by Binder, Smith and Holman. Bradford and Pickel and Blandford, in their CRIER (Cosmic-Ray-Induced-Error-Rate) analysis code introduced the basic Rectangular ParallelePiped (RPP) method for error rate calculations. For the radiation environment at the part, both made use of the Cosmic Ray LET (Linear Energy Transfer) spectra calculated by Heinrich for various absorber Depths. A more detailed model for the space radiation environment within spacecraft was developed by Adams and co-workers. This model, together with a reformulation of the RPP method published by Pickel and Blandford, was used to create the CR ME (Cosmic Ray Effects on Micro-Electronics) code. About the same time Shapiro wrote the CRUP (Cosmic Ray Upset Program) based on the RPP method published by Bradford. It was the first code to specifically take into account charge collection from outside the depletion region due to deformation of the electric field caused by the incident cosmic ray. Other early rate prediction methods and codes include the Single Event Figure of Merit, NOVICE, the Space Radiation code and the effective flux method of Binder which is the basis of the SEFA (Scott Effective Flux Approximation) model. By the early 1990s it was becoming clear that CREME and the other early models needed Revision. This revision, CREME96, was completed and released as a WWW-based tool, one of the first of its kind. The revisions in CREME96 included improved environmental models and improved models for calculating single event effects. The need for a revision of CREME also stimulated the development of the CHIME (CRRES/SPACERAD Heavy Ion Model of the Environment) and MACREE (Modeling and Analysis of Cosmic Ray Effects in Electronics). The Single Event Figure of Merit method was also revised to use the solar minimum galactic cosmic ray spectrum and extended to circular orbits down to 200 km at any inclination. More recently a series of commercial codes was developed by TRAD (Test & Radiations) which includes the OMERE code which calculates single event effects. There are other error rate prediction methods which use Monte Carlo techniques. In this chapter the analytic methods for estimating the environment within spacecraft will be discussed.