The TRADEX radar at the Kwajalein Missile Range (KMR) introduces a new debris characteriation capability to the space surveillance commuity. A single-sensor debris cataloging capability for small uncorrelated targets (UCTs) was recently demonstrated in a TRADEX autonomous catalog maintenance (TRACM) experiment. The experiment validated a number of new TRADEX space surveillance tools and integrated them with the unique Stare and Chase mode in which TRADEX finds and tracks small UCTs (objects less than 10 cm in diameter) at a rate of one new object per hour. Orbital elements for Stare and Chase acquisitions were determined from recorded track data and were of sufficient quality to permit re-acquisition by TRaDEX on subsequent passes. The experimental procedure is described and results are summarized in terms of successful re-tracks for objects maintained in a local proof-of-concept catalog. The catalog was established during the second half of a two month effort and individual debris objects were cataloged for up to 26 days. Comparison of TRADEX re-track performance for UCTs versus re-track performance for objects already cataloged by the space surveillance network (SSN) are also indicated. Conclusions and recommendations for further work are included.

There are many advantages to supplementing ground based radar debris detection systems with optical systems. For example: objects with a low radar signature can still be optically bright (and vice versa); in the field of space debris optical detection is less sensitive to range; the minimum detectable debris size for a given range is less than that for radar. Destructive debris can be as small as 1 cm, so any improvement in detection sensitivity towards this standard is important. To improve the accuracy of debris orbital elements, a real-time detection system might be preferable in contrast to one in which images are stored for post-observation ('daytime') analysis. This is because more than one telescope is needed to lengthen the observing baseline and so increase the detected fraction of the debris orbit. Therefore, any software based at one telescope that recognizes debris in its field of view, produces a first approximation of its orbit elements and alerts extra telescopes along track, must process its data quickly, and preferably during the same pass. A prototype of such a software system under development for use with a CCD camera at the Royal Greenwich Observatory's satellite laser ranger at Herstmonceux, East Sussex, England, is outlined. The methods which the detection algorithm employs to handle data from the camera system are described, along with the limitations that the hardware and processing time impose on the physical nature of the problem.

The Air Force Maui Optical Station (AMOS) has demonstrated follow-up astrometry and photometry for near-earth objects (NEOs), with results published in the Minor Planet Circulars. Although this information is important for the cataloging of all NEOs, it does not provide all of the data needed to assess the potential hazard posed by these objects, i.e. composition, size, shape, and dynamics. AMOS has increased its capability by adding a six position filter wheel (in conjunction with Phillip's Laboratory's Geophysics Directorate and the University of Arizona), for use on the Jet Propulsion Laboratory's CCD camera mounted on the AMOS 1.2 meter telescope. The paper provides the rationale for three-color photometry for determination of NEO characteristics, as well as preliminary results of the observations of several NEOs an main-belt asteroids. It also discusses the design of a photo-polarimeter, to be built in the near future, which will add more capability to AMOS, and determination of albedo and size of NEOs of particular interest to both the scientific and government communities.

Observational data on luminosity of bolides observed from ground-based stations as well as from satellite systems can be converted into masses or sizes of bodies assuming we know the luminous efficiency of the atmospheric interaction process. Problems of luminous efficiency were very much simplified until recently. Traditional approach assigned just a single value for each velocity and the calibration was derived only from early experiments with artificial meteors produced by masses of the order of grams. A recent analysis of the Lost City fireball enabled a precise and reliable determination of masses from the motion of the body and revealed that the differential luminous efficiencies for bodies in a mass ranges of hundred kilograms are out 10 times larger than the traditionally used values. This paper presents results on luminous efficiencies from detailed analysis of 29 Prairie Network (PN) bolides brighter than magnitude minus 10. This analysis is based on independent solutions for motion of the body, combines them with the experimental results for gram masses and applies them to satellite observed bolides. Masses resulting from such dynamic solutions are compared to radiated energies and the luminous efficiencies are computed by two different approaches: (1) total radiated energy is compared to the initial kinetic energy of the body resulting in total luminous efficiency (The only concept used on space observed bolides so far); (2) time change of kinetic energy of ablated mass is compared to power of radiation for each time-mark on the photographic record resulting in differential luminous efficiency (the traditional quantity used in meteor physics). Within a factor of two, the total luminous efficiency is identical to the differential one for majority of the 29 PN bolides. Some values of the total luminous efficiency are smaller than the differential ones due to velocity dependence and due to a significant part of kinetic energy going into changing the momentum of the body. Some values of the total luminous efficiency are greater than the differential ones due to sudden release of radiation energy at discrete points (bolides with large flares of many stellar magnitudes over the smooth light curve). Values of the differential luminous efficiencies for 29 PN bolides were derived altogether at 1324 different points of their trajectories. Correlations of these values with height, velocity, deceleration, and brightness were used to get correction factors for the actual bulk density and shape of the body (comparing these correlations with data on the Lost City bolide). These corrections were applied to the values of the total luminous efficiencies. The resulting total luminous efficiencies mostly depend on total radiated energy and, to a lesser extent, also on velocity. No reliable correlations with other parameters were found. Each bolide from these 29 behaves as an independent individual from the point of view of radiation efficiency in analogical way to ablation efficiency (Figure 8). There is no statistical correlation between ablation and radiation efficiency, but we can find groups of similar behavior: the most frequent are poor 'ablators' an good radiators (13 cases); the second in importance are good 'ablators' and poor radiators (9 cases). It is highly probable that this individualistic behavior is also valid for the much brighter bolides observed by DOD satellites. The average total luminous efficiency for the 29 PN bolides resulted as (tau) _t equals 1.36% of total kinetic energy amounting to an average of 5 multiplied by 10⁷ J of total radiated energy, which corresponds to the average initial mass of 15 kg for this sample of 29 PN bolides. These values were compared with experimental data on gram size meteoroids. Total luminous efficiencies and their expected errors for radiated energies in the range 10⁹ to 10¹³ were predicted (Table 5). Resulting total luminous efficiencies were applied to data published on 21 very bright bolides observed by sensors on DOD satellites and their masses were determined (Table 6).

During the early morning of November 21, 1995, a fireball as bright as the full moon entered the atmosphere over southeastern Colorado and initially produced audible sonic boom reports from Texas to Wyoming. The event was detected locally by a security video camera which showed the reflection of the fireball event on the hood of a truck. The camera also recorded tree shadows cast by the light of the fireball. This recording includes the audio signal of a strong double boom as well. Subsequent investigation of the array near Los Alamos, New Mexico operated by the Los Alamos National Laboratory as part of its commitment to the Comprehensive Test Ban Treaty negotiations, showed the presence of an infrasonic signal from the proper direction at about the correct time for this fireball. The Los Alamos array is a four-element infrasonic system in near-continuous operation on the laboratory property. The nominal spacing between the array elements is 212 m. The basic sensor is a Globe Universal Sciences Model 100C microphone whose response is flat from about 0.1 to 300 Hz (which we filter at the high frequency end to be limited to 20 Hz). Each low frequency microphone is connected to a set of twelve porous hoses to reduce wind noise. The characteristics of the observed signal include the onset arrival time of 0939:20 UT (0239:20 MST), with a maximum timing uncertainty of plus or minus 2 minutes, the signal onset time delay from the appearance of the fireball of 21 minutes, 20 seconds, total signal duration of 2 minutes 10 seconds, the source location determined to be toward 31 degrees from true north, the horizontal trace velocity of 429 m/sec, the signal velocity of 0.29 plus or minus 0.03 km/sec, assuming a 375 km horizontal range to the fireball, the dominant signal frequency content of 0.25 to 0.84 Hz (analyzed in the frequency interval from 0.2 to 2.0 Hz), the maximum signal cross-correlation of 0.97 and the maximum signal amplitude of 2.0 plus or minus 0.1 microbars. Also, on the basis of the signal period at maximum amplitude, we estimate a probable source energy for this event of between 10 to 100 tons of TNT (53.0 tons nominal).

A new European Space Agency (ESA) flight instrument attached to the exterior of the MIR Space Station is providing a better understanding of the effects of the space environment. The instrument was designed to measure, real time, the impacts and trajectory of hypervelocity particles, the atomic oxygen flux and contamination deposition/effects during the course of the mission. The ESA mission, EuroMir'95, began in September 1995 and was completed in March 1996. Active data from the momentum detectors have reconfirmed the existence of an orbital debris cloud. The mission also allowed for an EVA which returned passive materials to Earth for subsequent laboratory analyses. The early results of this experiment suggest the existence of one reasonable size cloud of small size debris particles with momenta in the range of 4E-11 kg-m/s to 5E-10 kg-m/s. These data are considered quite germane due to the similarity in orbital altitude and inclination of the Mir and Alpha Space Stations.

On retrieval from low Earth orbit (LEO), the solar arrays from the European retrievable carrier (EuReCa) and one solar array wing of the Hubble Space Telescope (HST) were inspected for micrometeoroid and space debris impact damage. Seven-hundred-three and eight-hundred-fourteen impact sites respectively were analyzed in detail. Interpretation of particle parameters from this large data set can yield a useful measurement of the micrometeoroid and debris flux in LEO. Due to similar orbital parameters, this then provides a flux measurement complimentary to those measured by the timeband capture cell experiment (TiCCE) of the EuReCa spacecraft and detectors and exposed surfaces on the Long Duration Exposure Facility (LDEF). To allow comparison of data from the space retrieved solar cells to previous data, solar cell material, glass and aluminum targets were impacted for inter-calibration using the University of Kent's Light Gas Gun (LGG). An average of 20 impact sites on each target material per shot were measured. It was decided to consider only non-perforations of the 150 micron thick CMX cover glass, common to both EuReCa and HST solar cells. Trends in crater morphology of the laboratory impacts are discussed and compared to those from space impacts. The effects of impact angle and crater scaling with particle size are investigated and a conversion from appropriate solar cell crater parameters to the ballistic limit in aluminum is presented.

This paper proposes a methodology for combining optical and radar observations to determine material types of debris objects. Determining the material types of debris objects will in turn allow more accurate estimates of the sizes and masses of debris objects. In order to estimate the capabilities of this technique, simulated optical and radar data is used to determine the material types and object sizes of simulated debris objects. Laboratory test data measuring the radar and optical responses of several common spacecraft materials are presented and are used in the simulations. Albedo data and a set of radar response curves for several materials commonly used in spacecraft construction are presented. The capabilities of this method are discussed, and conclusions are reached.

Determining the hazards from debris-generating events is a design and safety consideration for a number of space systems, both currently operating and planned. To meet these and other requirements, the United States Air Force (USAF) Phillips Laboratory (PL) Space Debris Research Program has developed a simulation software package called the Debris Analysis Workstation (DAW). This software provides an analysis capability for assessing a wide variety of debris hazards. DAW integrates several component debris analysis models and data visualization tools into a single analysis platform that meets the needs for Department of Defense space debris analysis, and is both user friendly and modular. This allows for studies to be performed expeditiously by analysts who are not debris experts. The current version of DAW includes models for spacecraft breakup, debris orbital lifetime, collision hazard risk assessment, and collision dispersion, as well as a satellite catalog database manager, a drag inclusive propagator, a graphical user interface, and data visualization routines. Together they provide capabilities to conduct several types of analyses, ranging from range safety assessments to satellite constellation risk assessment. Work is progressing to add new capabilities with the incorporation of additional models and improved designs. The existing tools are in their initial integrated form, but the 'glue' that will ultimately bring them together into an integrated system is an object oriented language layer scheduled to be added soon. Other candidate component models under consideration for incorporation include additional orbital propagators, error estimation routines, other dispersion models, and other breakup models. At present, DAW resides on a SUN^R workstation, although future versions could be tailored for other platforms, depending on the need.

There has been some recent concern in the orbital debris scientific community over whether the Poisson probability density function is valid for the present-day studies involving larger debris fluxes, longer duration missions, and larger cross sectional areas. This paper presents the mathematical definitions for Poisson and its more physically founded cousin, the binomial distribution, and discusses the relationship between them. The debris impact problem is then formulated using both distributions, and a quantitative comparison is presented along with several examples. It is concluded that although the orbital debris collision problem is inherently binomial, one often does not have the necessary knowledge to accurately formulate the binomial expression. Furthermore, Poisson provides probability estimates that, for a wide range of studies, approximates those of the binomial very well.

The population of dangerous orbital debris particles in low Earth orbit is growing, resulting in a need for improved risk assessment and risk management of critical space station elements from this hypervelocity impact threat. The effects of micrometeoroid and orbital debris (M/OD) penetration on space station equipment can very from a local damage problem to a possible catastrophic system failure (i.e., crew or station loss). Given this possibility, a preliminary study was undertaken by Meyer Analytics and NASA-MSFC to quantify and reduce the likelihood of catastrophic failure following orbital debris penetration of two Russian supplied modules: the FGB Energy Block module and the SPP-1 gyrodyne module. Each of these modules contains stored energy in the form of chemical propellants (UDMH), high pressure tanks, or kinetic energy (gyrodynes) that may release energy and propagate damage to the rest of the space station if impacted by a sufficiently energetic orbital debris particle. The study recommended design improvements to lower these probabilities, including spall blankets around the gyrodynes to lower the probability of gyrodyne penetration and fragment damage given gyrodyne rupture.

This report describes an investigation with the objective of substantiating the current hypothesis that describes the source of continually recurring space borne optical sensor readings of unknown origin received from several monitoring satellites. These unidentified optical signals have several characteristics such as long duration, short rise time, and magnitude, which do not correspond to expected signals. Up to now, the limited qualitative evidence indicates the possibility of micro-particle impact induced signals. A quantitative investigation is presented from laboratory experiments and numerical analysis which further support this hypothesis. Sources that can be definitely eliminated include signals from lightning, solar compensation, platform motion, or NUDET. Additionally, it can be concluded from several pieces of evidence that the source is local to the satellite. We can also eliminate the possibility of a hardware anomaly of a specific sensor design since data from 1988 to the present are being detected by each of the different types of optical sensors used (3 satellite platforms and 5 different optical sensor designs are involved). Therefore, it is believed that these signals are a consequence of some real, external physical phenomenon and not just a consequence of some anomalous sensor behavior. The current evidence is not a sufficient basis for a complete description of the phenomenology and needs to be studied in greater detail to verify whether or not hypervelocity particle impact is indeed the source of the optical triggers described above, and if so, to work out the phenomenology in quantifiable terms. To substantiate the hypothesis of a HV particle impact source for these signals, the physical phenomenon was reconstructed both experimentally and computationally. Detailed numerical calculations representing the interaction of a specific particle size and velocity were conducted. The intensity of the impact flash is inferred based on temperatures using a black body radiation assumption. Laboratory experiments consisted of HV impacts of iron particles onto a lens specimen and sunshade specimen using a 2 MV Van De Graaf accelerator. Results of the experimental data obtained using a photomultiplier are compared to the numerical calculations and with existing data.

Shadowgraphic holography allows imaging of small particles over a 180 degree field of view and with a large depth of field. The cylindrical holography technique developed by Hough and Gustafson has been modified to allow holograms of small hypervelocity impact generated simulated space debris particles to be successfully made by using a subnanosecond laser pulse length. The use of very short (135 ps) laser pulses with a corresponding short coherence length (4 cm) frees the motion of small high speed particles. With this system, shadowgraphic holograms of aluminum projectiles impacting aluminum and graphic epoxy plates have been achieved at hypervelocity. Results of these tests as well as low speed proof of concept tests are presented.

To protect spacecraft from orbital debris requires a basic understanding of the processes involved in hypervelocity impacts and characterization of detectors to measure the space environment. Both require a source of well characterized hypervelocity particles. Electrostatic acceleration of charged microspheres provides such a source. Techniques refined at the Los Alamos National Laboratory provided information on hypervelocity impacts of particles of known mass and velocity ranging from 20 - 1000 nm diameter and 1 - 100 km/s. A Van De Graaff generator operating at 6 million volts was used to accelerate individual carbonyl iron microspheres produced by a specially designed particle source. Standard electrostatic lenses and steering were used to control the particle's flight path. Charge sensitive pickoff tubes measured the particle charge and velocity in-flight without disturbing the particle. This information coupled with the measured Van De Graaff terminal voltage allowed calculation of the particle energy, mass, momentum, and using an assumed density the size. Particles with the desired parameters were then electrostatically directed to a target chamber. Targets used in our experiments included cratering and foil puncture targets, microphone momentum enhancement detectors, triboluminescent detectors, and 'splash' charge detectors. In addition the system has been used to rapidly characterize size distributions of conductive plastic particles and potentially provide a method of easily sorting microscopic particles by size.

Long duration spacecraft in low earth orbit, such as the International Space Station Alpha (ISSA), are highly susceptible to hypervelocity impacts by pieces of debris from past earth-orbiting missions. With this increased likelihood of debris impact over time comes a responsibility on the part of a spacecraft design engineer to quantify, and subsequently reduce, the hazardous effects on the spacecraft and its crew should a penetration occur. Among the various hazards that accompany the penetration of a pressurized manned spacecraft module are critical crack propagation in the module wall (i.e. the so-called 'unzipping' of the module) and depressurization-related phenomena, such as crew hypoxia and uncontrolled thrust due to air rushing out of the module wall hole. These phenomena are directly related to the hole size and crack lengths that result in a spacecraft wall following a penetration. As a result, accurate models for hole size and crack size resulting in thin-walled spacecraft structures following hypervelocity impact penetration are required for accurate spacecraft quantitative risk assessments. This paper presents the results of a study whose objectives were to develop semi- empirical models of hole size and tip-to-tip crack length for some of the multi-wall shielding systems being developed for ISSA. The empirical models were developed using light gas gun test data at impact velocities around 6.5 km/sec, and inhibited shaped charge test data at an impact velocity of 11.3 km/sec. The significance of the work performed is that these models can be incorporated directly into a survivability analysis to determine whether or not module unzipping would occur under a specific set of impact conditions. In addition, the prediction of hole size can be used as part of a survivability analysis to determine the time available for module evacuation prior to the onset of incapacitation due to air loss.

The IRIDIUM Program represents the most exciting prospect in space operations in the world today. It represents one of the best commercial opportunities as well as one of the biggest engineering challenges ever. This paper describes the trend-setting approach Motorola has taken in 'designing in' orbital debris mitigation features while keeping associated costs virtually negligible. It delineates the attributes of a comprehensive program instituted during the design phase, maintained through the development phase, and which will be continued through the operations phase. The challenges are being met head-on and Motorola is aggressively seizing the opportunity to set new standards in debris protection mitigation techniques.

The development of debris environment analysis tools and correlation of space weather with satellite anomalies have progressed nearly independently during the space age. This paper proposes how these two fields may be merged to provide greater insights into the operation of satellites in the dynamic space environment. The combination of these two applications will provide a venue to corroborate the impact of debris on spacecraft while identifying the cause of previously 'unknown' anomalies. A conceptual approach as how this may be accomplished is presented using events to highlight the efficacy of this effort.

This paper introduces a strategy for minimizing the growth of orbital debris by controlling the collision hazard that it represents to operational satellites. Applying the theory of Kessler, an analytic tool can be developed to set target survivability levels for operational satellites, to determine the individual contributions of different missions to the debris hazard, and to manage the debris environment by assessing the efficacy of different mitigation measures.

As the number of Earth orbiting objects increases so does the potential for generating orbital debris with the consequent increase in the likelihood of impacting and damaging operating satellites. Various debris remediation approaches are being considered that encompass both in-orbit and return-to-Earth schema and have varying degrees of operations, cost, international competitiveness, and safety implications. Because of the diversity of issues, concerns and long-term impacts, there is a clear need for the setting of government policies that will lead to an orderly abatement of the potential orbital debris hazards. These policies may require the establishment of a supportive regulatory regime. The Department of Transportation is likely to have regulatory responsibilities relating to orbital debris stemming from its charge to protect the public health and safety, safety of property, and national security interests and foreign policy interests of the United States. This paper describes DOT's potential regulatory role relating to orbital debris remediation, the myriad of issues concerning the need for establishing government policies relating to orbital debris remediation and their regulatory implications, the proposed technological solutions and their economic and safety implications. Particular emphasis is placed upon addressing cost-effectiveness and economic analyses as they relate to economic impact analysis in support of regulatory impact analysis.

Space is traditionally viewed as boundless and vast and, therefore, relatively little attention has been paid to the physical locations of operational satellites for the purpose of avoiding collisions. The advent of large satellite constellations has focused new light on the issue of physical interference between objects in space. Concentrating on U.S.-licensed medium an low Earth orbit (collectively referred to as LEO) communications constellations, this article examines the need for 'traffic regulation' in the LEO environment. It notes the emergence of voluntary coordination and disposal practices and explores whether additional government oversight and guidelines are necessary. Concluding in favor of some government guidelines, the article discusses who can impose them.

Following release of the 1989 'Report on Orbital Debris' by the Interagency Group (Space) for the National Security Council, NASA undertook a series of extensive bilateral discussions with the major spacefaring nations on the topic of orbital debris. These discussions led to a greater understanding of both the cause and the effect of orbital debris. As a result of these discussions, the major spacefaring nations have taken definitive steps to redesign their launch vehicles and spacecraft so as to mitigate the production of orbital debris. In 1993 the National Aeronautics and Space Administration (NASA), the European Space Agency (ESA), and Japan formed a multilateral Inter- Agency Orbital Debris Coordination Committee (IADC). Since that time the Russian Space Agency (RSA), the Chines National Space Agency (CNSA), the French National Space Agency (CNES), the British National Space Agency (BNSA), and the Indian Space Agency (ISRO) have jointed the IADC. In 1994 orbital debris discussions began in the United Nations under the auspices of the Scientific and Technical Subcommittee of the Committee on the Peaceful Uses of Outer Space (UNCOPUOS). In 1995 UNCOPUOS adopted a multi-year program for studying orbital debris. In 1993 the White House Office of Science and Technology Policy (OSTP) and the National Security Council (NSC) undertook an interagency review to revise and update the 1989 'Report on Orbital Debris.' In November 1995 Dr. John H. Gibbons, the Assistant to the President for Science and Technology, released the 'Interagency Report on Orbital Debris -- 1995.'

Natural coordinates method for solution of some dynamic problems is proposed. Advantages of the method are illustrated by examples of the following well-known problems connected with (1) the motion of two non interacting particles in absence of external forces; (2) the motion of a charged particle in the magnetic field of a monopole; (3) the motion of a particle in Coulomb or gravitational fields. Analytical solutions of the problems have been determined. Classification of motions and according graphic illustrations have been done.

Set of orbital elements of a celestial body naturally originating from dynamic problem solution is proposed. Orbital elements considered are 4 spherical coordinates of radius-vector and body velocity vector for the moment of passing a pericenter and time of this passing. Dependencies, connecting proposed orbital elements and conventional orbital elements of celestial mechanics have been determined. These enable one to perform direct and inverse coordinate transformations. Advantages of usage of natural orbital elements have been illustrated by example in which probability distributions of natural orbital elements for comets have been determined. Analysis of these probabilities enables one to make some suggestions on comets origin and to classify them accordingly. Another advantage of usage of natural orbital elements consists in determination of analytical dependence between a perihelion velocity and a perihelion distance.

This paper addresses the process by which the range safety of intercept tests are assessed. The methodology outlines procedures for consideration of the safety of space, airborne, and land fetures. A special emphasis is placed on the role of breakup models on the utility of the final result. Issues covered include determining hazards vs. avoidance zones, fragment propagation, the impact of breakup model fidelity or resulting simulations, and lethality limits for space/air/ground.