The Mars Microprobe Project provides a unique opportunity to validate instrumentation technologies for performing relevant science experiments on the Martian surface and at a depth of up to two meters. The instrumentation package of the Mars Microprobe includes a meteorological instrument for measuring atmospheric pressure and temperature, accelerometers to measure the atmospheric drag on the probe aeroshell during descent, accelerometers for measuring the forces on the probe at impact, temperature sensors in the penetrator to measure the soil thermal conductivity at depth, and an experiment to collect a small sample of the Martian regolith and determine if water is present in the sample. These instruments have been designed to withstand the deceleration forces resulting from the impact and penetration of the probe into the Martian surface and to function in the harsh temperature environment beneath the surface. Results from the design and environmental test of these instruments are presented.

The proposed Space Interferometry Mission (SIM) spacecraft carries high resolution stellar interferometers for micro-arc- second accuracy astrometric measurements. These stellar interferometers require picometer accuracy one dimensional metrology gauges, surface metrology gauges and 3-dimensional metrology gauges. The absolute metrology gauges required by these interferometers can be considerably less accurate due to the careful design of the astrometric interferometers on the spacecraft. Open-faced, hollow corner cube retro-reflectors are used as fiducials in the one-dimensional relative and absolute metrology gauges and the 3-dimensional metrology gauge. The diffraction caused by the assembly and the component defects of these hollow retro-reflectors affects the accuracy of these metrology gauges. A simulation quantifying some of the effects of the component and assembly defects of hollow retro-reflectors on the accuracy of a picometer linear metrology gauge is presented. An auto-aligning, 3-dimensional metrology gauge constructed using the sub-picometer linear metrology gauges was described in earlier papers. The functioning automatic alignment and the sub-nanometer, in-air tracking results from this 3-dimensional metrology gauge are presented.

Creative new optical designs for coronagraphs which use only reflecting elements are extremely well suited for planetary studies which usually require detection of large, faint, tenuous sources about bright central planets (themselves worthy of study). These new coronagraphic designs not only allow the observation of extended atmospheres and coronae, they also allow critical observations of the central planet at the same time with instruments optimized for different wavelengths. The new coronagraphic systems can be more easily accommodated within the envelope of launch vehicle capabilities available today than can older, slower systems, and they permit simple spacecraft designs which reduce weight, power, and cost. They possess inherently higher end-to-end optical efficiencies. The very modest fluxes associated with many extended sources in the solar system, however, require state-of-the-art fabrication techniques, and place new demands on focal plane instrumentation. We focus here on an instrument designed to observe the lunar atmosphere. Also considered are several archetypical problems, including the study of the neutral cloud an ionized torus associated with Jupiter's moon Io and of comets.

The recent availability of highly integrated CMOS image sensing arrays at low cost has made the goal of an affordable, low power, low weight attitude control sensor attainable. One such device is incorporated into the MSM-100 from Santa Barbara Sensor Technologies and includes a 160 by 160 sensing array, clock driving circuitry and an eight bit A/D converter. Output may be selected as either the analog pixel signal level or digitized data in either parallel or serial format. This sensor was coupled with an 8 mm f/1.3 lens and tested for star image sensitivity by Santa Barbara Sensor Technologies. Signal data were collected from stars of magnitude 1.5 and brighter in a 12 degree square field of view at a rate of 1.2 frames per second. The sensitivity limit in the present configuration is a star of magnitude 2.0 at a S/N ratio of 10.

Our previous studies reported on the development of directional sensors for gamma and x-rays that have an angular resolution of one second of arc and are limited only by the accuracy of our measurement of the angle. Those sensors produced response functions that looked like step functions when the detectors were rotated in the fields of gamma ray or x-ray photons generated by point sources. The sensors located point sources accurately at intensities as low as 10 photons per cm²-s in the presence of much higher intensities from other sources. Their unique ability is that they can distinguish between photons emitted by a point source that did not interact on their way from the source to the detector from those that did interact. This paper reports on the design and fabrication of a new directional scintillation sensor with a sensitivity 45X greater than previous designs. The results show that the new sensor can be used in gamma and x-ray astronomy.

Two spacecraft built by the Johns Hopkins University Applied Physics Laboratory carried a total of six spectrographs into space early in 1996; five were on the Midcourse Space Experiment and one was on the Near Earth Asteroid Rendezvous Spacecraft (NEAR). Flight spectrographs have also been delivered for the Defense Meteorological Satellite Program and the Global Ultraviolet Imaging spectrograph (GUVI) is currently being built for the TIMED satellite. These eight instruments, while on four different spacecraft and working in different spectral ranges extending from 115 nm out to 2700 nm, were based on only two optical designs. The five spectrographs on the MSX used the same design and the DMSP, NEAR, and GUVI used an identical optical layout even though the wavelength range changed from the vacuum ultraviolet to the infrared. This paper reviews the performance parameters and optical designs of these instruments and shows how minor optical changes made it possible to use the identical mechanical design for a range of applications. Rather than create a unique design for each application, modification of an existing design resulted in a saving of cost, time, and testing. In spite of the plagiarized designs each instrument is fully capable of meeting all of its scientific objectives. Designs for future missions follow the same concept; a basic optical layout is prepared and is then modified to meet the field-of-view, spectral range, and capability requirements of a mission. Some proposed concepts for future spacecraft are also described.

In this paper some characteristics of photothermoplastic carriers (PTPC) are presented, and also the parameters of the vision slit apparatus, created on base of PTPC. The achieved parameters of the relief-phase recorded images in the process of work, constitute rather perspective and are able, on our opinion, to attract the explicators of future cosmic apparatus.

Proposals for determining the gravitational constant G in space are compared and contrasted. We find that only three proposals have carefully treated the many physical processes in the space environment which might potentially vitiate the sought-for accuracy. The capability of the proposed missions to make other gravitational measurements and tests is described.

A proposed space-based test of gravitational theory requires unique performance for thermometry and ranging instrumentation. The experiment involves a cylindrical test chamber in which two free-floating spherical test bodies are located. The test bodies are spheres which move relative to each other. The direction and rate of motion depend on the relative masses and orbit parameters mediated by the force of gravity. The experiment will determine Newton's gravitational constant, G; its time dependence, as well as investigate the equivalence principle, the inverse square law, and post- Einsteinian effects. The absolute value of the temperature at which the experiment is performed is not critical and may range anywhere from approximately 70 to 100 K. However, the experimental design calls for a temperature uniformity of approximately 0.001 K throughout the test volume. This is necessary in order to prevent radiation pressure gradients from perturbing the test masses. Consequently, a method is needed for verifying and establishing this test condition. The presentation is an assessment of the utility of phosphor-based thermometry for this application and a description of feasibility experiments. Phosphor thermometry is well suited for resolving minute temperature differences. The first tests in our lab have indicated the feasibility of achieving this desired temperature resolution.

In this paper, we report on the mission study and laboratory development for ASTROD. ASTROD mission concept is to use drag- free spacecraft in solar orbits together with a constellation of Earth orbiting satellites to provide high-precision measurement of relativistic effects, better determination of the orbits of major asteroids, improvement in the measurement of G, measurement of solar angular momentum via Lense-Thirring effect and the detection of low-frequency gravitational waves and solar oscillations in a single mission. We present our progress in the mission study on the various aspects for measuring these quantities. As to the laboratory development, after reviewing briefly our progress in log fiber-linked heterodyne interferometer and tunable fiber directional coupler, we present our progress in weak light optical phase locking, fiber delay-line and laser and time metrology.

A program for investigating planetary rings must include purposeful experiments to explore the superconductive material state of the rings. Our hypothesis of the superconductive material state of planetary rings makes it possible to add classical theories of planetary rings with a nonconflicting superdiamagnetic model. It solves many problems of planetary rings by considering the phenomena peculiar to superconductors during interaction with the magnetic field.

NASA Management Instruction (NMI) 1700.8 directs each project office to limit orbital debris generation if this action is cost-effective and consistent with achieving mission objectives. To implement this policy, the NASA Office of Safety and Mission Assurance, the sponsor of NMI 1700.8, tasked NASA Johnson Space Center (JSC) to develop the NASA Safety Standard 1740.14: Guidelines and Assessment Procedures for Limiting Orbital Debris, August 1995. To mitigate the accumulation of mass in Earth orbit, NSS 1740.14 addresses the issues of postmission disposal of spacecraft and upper stages. According to the guidelines, these systems in general should be left in an orbit in which, using conservative projections for solar activity, atmospheric drag and gravitational perturbations will limit the lifetime in low Earth orbit (LEO) to no longer than 25 years after completion of mission. Consequently, JSC undertook a series of studies to investigate the most efficient and cost effective options for reducing orbit lifetime. In this paper we present an overview of the various options and give hints for the choice of the option best suited for specific mission types, e.g., depending on initial orbit, existing propulsion systems, existing electrical power level, electrical power and attitude control lifetime, and acceptable maneuver time and mass penalties.

During the period February 1995 - May 1997, the U.S. Space Shuttle visited the Russian Mir space station on one close rendezvous and six docking missions. A detailed test objective (DTO-1118) called for extensive photographic and video imagery of the Mir complex for several purposes, including to assess the overall condition of the station and to study the effects of the space environment on a long-duration orbiting platform. Thousands of photographs of Mir from 35 mm Nikon and 70 mm Hasselblad cameras were taken, and more than one hundred hours of video from several cameras located in the Space Shuttle cargo bay were collected. A review of these photographic data has revealed evidence of numerous small particle impacts. This paper describes the photographic analysis effort at the NASA Johnson Space Center with an emphasis on Mir particulate damage assessment. Sample photographs depicting impact effects are provided. A preliminary attempt is made to compare the observed impacts with predictions based on environmental models. A comprehensive comparison between data and the models is hampered by the lack of a complete historical record of the attitude of the Mir space station and by an inability to distinguish between orbital debris and meteoroid impacts from only remote photographic evidence.

Improvements in hardware and processing techniques for the passive detection of infrasonic sound waves provide capabilities for monitoring space debris and meteor impacts. A review of past detection of meteors indicates possible uses for infrasonic observatories. We also document signals from bolides that we recorded on 3 - 4 October 1996, illustrating long-range tracking capabilities and review short-range detections and the observation of a space shuttle reentry. We discuss possible roles of acoustic observing systems for space debris and meteor detection.

The purpose of this report is to present the engineering design of a spaceborne micro particle impact detector (MPID) experiment. This experiment is manifested on a Phillips Laboratory spacecraft called MightySat I scheduled for launch in July 1998. A follow-on report will present the resulting particle impact data. The objective of this experiment is to measure direction and time of impact of spaceborne micron size particles with time of impact resolution of 0.1 seconds. The primary element in this experiment consists of two metal- oxide-semiconductor (MOS) discharge capacitor detectors that discharge upon hypervelocity particle impact. The detectors were developed by Prof. J. J. Wortman from North Carolina State University. Each MOS particle detector is 3 in by 1-1/2 in and approximately 0.013 in thick. Each particle detector is bonded to a detector assembly that is in turn mechanically fastened to the external bottom plate of the MightySat I spacecraft. The detector assembly and associated electronics weigh less than 0.4 lb and have a total impact detection area of 3.7 in². Each particle impact causes an impact event record to be stored in the spacecraft control unit for later downlink. Each impact event record will store time of impact and output from two coarse sun sensors. Data from the coarse sun sensors is used to help determine attitude of the spacecraft. The Phillips Laboratory MightySat I spacecraft, developed largely by CTA Space Systems in McLean, Virginia, designed for ejection from the Space Shuttle is a 6-sided composite structure, 20.5 in (height) by 19.0 in (diameter), 150 lb., and spin stabilized with 5 degree attitude knowledge. The MightySat I spacecraft is scheduled for orbit injection using a standard hitchhiker ejection system from space shuttle flight STS-88. (Ref. 1)

Data on the influx of interplanetary bodies onto the Earth in the size range from millimeters to kilometers are derived from observations of meteors and bolides of smaller sizes and from Spacewatch discoveries of asteroids of larger sizes. Special attention is given to a new calibration of meteoroid masses. The latest published interpretation and calibration of the rate of discoveries of the Spacewatch asteroids is incorporated into the final flux curve. In the mass range from 10 exp 4 to 10 exp 7 kg, infrasound data are also discussed in some detail, and the resulting cumulative flux is given.

To obtain photographic records of bright fireballs, many stations over a large territory, each with cameras covering the whole sky, are needed. The number of stations operating in Central Europe nowadays hovers around 40, and the average spacing between them is approximately 100 km. The whole system covers an area of about one million sq km. The most developed observational system of the European Network (EN) is in the Czech part, where each station is equipped with one fixed fish-eye camera. The extremely good optical quality of Zeiss Distagon objectives used in these cameras enables us to derive positions from one photograph of the whole sky hemisphere with a precision better than 1 arcmin. The German part is mostly equipped with less precise all-sky mirror cameras. The capabilities of this observational system are demonstrated on two very recent exceptional cases.

Decays of satellites 1991-005B and 1994-087B were photographed by fish-eye cameras of the European Fireball Network. A method of computation of the curved satellite trajectories from the photographs taking into account Earth rotation was developed. The position of decaying satellites in space was determined with the precision of about 100 m along a 1000-km long path. Velocity could be determined on a part of the path. The results are compared with trajectories predicted with a satellite tracking program.

Although a variety of orbital debris measurement data is available, all these data together do not characterize the orbital debris and meteoroid environment in a way that allows the direct estimation of potential hazards for active and planned space missions. This can only be done by modeling. The measurement data can be used for the evaluation of modeling results and for the calibration of the models themselves. In this paper it is shown how two-line element sets (TLE), radar cross-section data (RCS) and satellite catalog data are compared to current NASA breakup model results. It is shown that neither the assumption of a fixed lower trackable size threshold nor of completeness of the satellite catalog above a certain size are adequate for comparison purposes. A solution for this problem, i.e., a better way to handle the data, is presented. Furthermore a realistic picture of the growth and evolution of the total population in orbit is given.

The International Space Station (ISS) will be at risk from orbital debris and micrometeorite impact (i.e., an impact that penetrates a critical component, possibly leading to loss of life). In support of ISS, last year the authors examined a fundamental assumption upon which the modeling of risk is based; namely, the assertion that the orbital collision problem can be modeled using a Poisson distribution. The assumption was found to be appropriate based upon the Poisson's general use as an approximation for the binomial distribution and the fact that is it proper to physically model exposure to the orbital debris flux environment using the binomial. This paper examines another fundamental issue in the expression of risk posed to space structures: the methodology by which individual incremental collision probabilities are combined to express an overall collision probability. The specific situation of ISS in this regard is that the determination of the level of safety for ISS is made via a single overall expression of critical component penetration risk. This paper details the combinatorial mathematical methods for calculating and expressing individual component (or incremental) penetration risks, utilizing component risk probabilities to produce an overall station penetration risk probability, and calculating an expected probability of loss from estimates for the loss of life given a penetration. Additionally, the paper will examine whether the statistical Poissonian answer to the orbital collision problem can be favorably compared to the results of a Monte Carlo simulation.

Statistical methods for expressing the impact risk posed to space systems in general [and the International Space Station (ISS) in particular] by other resident space objects have been examined. One of the findings of this investigation is that there are legitimate physical modeling reasons for the common statistical expression of the collision risk. A combination of statistical methods and physical modeling is also used to express the impact risk posed by re-entering space systems to objects of interest (e.g., people and property) on Earth. One of the largest uncertainties in the expressing of this risk is the estimation of survivable material which survives reentry to impact Earth's surface. This point was recently demonstrated in dramatic fashion by the impact of an intact expendable launch vehicle (ELV) upper stage near a private residence in the continental United States. Since approximately half of the missions supporting ISS will utilize ELVs, it is appropriate to examine the methods used to estimate the amount and physical characteristics of ELV debris surviving reentry to impact Earth's surface. This paper examines reentry survivability estimation methodology, including the specific methodology used by Caiman Sciences' 'Survive' model. Comparison between empirical results (observations of objects which have been recovered on Earth after surviving reentry) and Survive estimates are presented for selected upper stage or spacecraft components and a Delta launch vehicle second stage.

Although the current hazard to most space activities from debris is low, growth in the amount of debris threatens to make some valuable orbital regions increasingly inhospitable to space operations over the next few decades.

During the evening of October 3, 1996, at least six bright fireballs were observed over the western United States with reports from California to Louisiana. The event over California produced tremendous sonic boom reports in the Los Angeles area. This event was also detected locally by 31 seismometers which are part of a network of seismic stations operated by the California Institute of Technology. Subsequent investigations of the data from the four infrasound arrays used by LANL (Los Alamo National Laboratory) and operated for the DOE (Department of Energy) as part of the CTBT Program (Comprehensive Test Ban Treaty) Research and Development program showed the presence of an infrasonic signal from the proper direction at the correct time for this bolide from two of our four arrays (Nevada Test Site; NTS and Pinedale, WY; PDL). Both the seismic and infrasound recordings indicated that an explosion occurred in the atmosphere, having its epicenter near Little Lake, Calif. for possible sources heights from 40 - 60 km. The infrasonic arrays are each composed of four elements, i.e., low frequency pressure sensors that are in near-continuous operation. The nominal spacing between elements is 150 - 200 m depending on the specific site. The basic sensor is a Globe Universal Sciences Model 100C microphone whose amplitude response is flat from 0.1 to 300 Hz. Each sensor is connected to 12 porous hoses which act to reduce wind noise. The signal characteristics, analyzed from 0.1 to 5.0 Hz, includes a total duration of 5 (NTS) to 20 minutes (PDL) for a source directed toward 230 - 240 degrees from true North. The signal trace velocities ranged from 300 - 360 m/sec with a signal velocity of 0.30 plus or minus 0.03 km/sec, implying a stratospheric (S type) ducted path (with a reflection altitude of from 40 - 60 km). The dominant signal frequency is from 0.20 to 0.80 Hz, with a peak near 0.2 to 0.25 Hz. These highly correlated signals had a maximum amplitude of 1.0 microbars (0.1 Pa) at PDL and 4.0 microbars (0.4 Pa) at NTS. Our analysis indicates that the bolide had a probable, maximum source energy in the range from 150 - 390 tons (TNT equivalent).

Solid-fuel rocket motors are well recognized as a source of numerous small-sized (10 micrometer or less) debris that are ejected at high velocities during the propellant burning process. Medium-sized (1 mm to 10 cm), low velocity versions of these metallic oxide or other combustion chamber debris have also been reported from static ground tests of solid-fuel motors. Field observations of a third component of the debris generated by solid-fuel rocket motor operation are presented in this paper. These are medium-sized debris that are expelled at low velocities through the rocket motor nozzles after the nominal cessation of propellant burning. These post-burnout debris, referred to as chuffing debris, may be a significant component of the orbital debris environment. Radar and optical measurements of these debris have been collected during numerous sub-orbital flight tests conducted over the past several years. The large database of such observations that has now been accumulated indicates that such post-burnout debris are a generic consequence of solid-fuel rocket motor operation. Selected portions of this database are reviewed, and a preliminary model of such medium-sized debris production is presented that is suitable for correlation with existing orbital debris observations and population models.

This is a concept study for a proposal to NASA/GSFC for a medium class Explorer Mission. It is designed to replace a prior SBIR Phase I design for NASA/MSFC for a Lunar far-UV survey telescope done in 1994 - 1995 for the Pathfinder Program (by the authors for I.S.E., under M. E. Nein, MSFC). A full investigation by project scientist D. L. White as to the most desirable mission science for a Lunar-based UV telescope, resulted in the decision to do a universal survey of the most interesting lines in the Lyman alpha forest, especially the O VI doublet lines around 103.2/103.8 nm. A telescope was designed by the authors incorporating a multiple instrument pod (MEDUSA), and a unique optical train featuring a selectable element secondary mirror module, with a special high resolution mode debuting a new optical design, all by chief optical engineer E. W. Cross. Special thanks go to chief spacecraft engineer T. L. Kessler for all packaging and integration of the telescope, its attendant systems, and the entire mission, including the launch interface and all presentations. In this incarnation, the basic concept has been converted by D. L. White into a free flyer designed for at least a LEO. In reconfiguring the original concept in the order to accomplish the original mission science goals, it has been necessary to take a fresh approach in order to fit the largest feasible Explorer Class Fairing (10L). In addition, the reconsideration of the mission science and the performance level available from the prior mission's optics, the authors decided to push the limits of the possible in the pursuit of excellence and choose two exceptional optical designs, augment them, and integrate them into the same limited envelope, while not sacrificing performance, communications, power, control, or serviceability. This we have kept close to focus throughout our pursuit of the mission science, which we hold foremost. We see a great need to bring the lessons learned at other portions of the spectrum to bear in this very densely populated, yet still not fully explored region. We wish to provide missions such as FUSE with the best survey data, to provide for the successor/s to FUSE, and to also do valuable science Our goal remains to provide an advancement to the field of F-UV astronomy by utilizing the best of the designs and engineering available to us, while incorporating on our team the direct participation of highly talented, as well as experienced professionals from different disciplines, while remaining open to the invaluable critiques and collaborations available in the community of which we are a part