The conception of ground-based optical observations of 'Space Debris'--pieces is described. This conception is based on the experience of photographic and TV observations of geostationary satellites and some theoretical investigations that have been carried out at the Institute for astronomy of Russian Academy of Sciences. The following problems are considered in this report: positional and photometric observations of Space Debris in GEO; the optimization of optoelectronic observations; the advantages of base-line optical observations. The secular evolution of Space Debris particles on high-latitude orbits due to solar radiation pressure is also described.

As part of Air Force Phillips Laboratory's Space Debris Program, MIT Lincoln Laboratory's Experimental Test System has conducted three separate types of surveys to detect small uncatalogued debris. In addition, follow-up metric and photometric observations have also been made. The goal of these observations is to characterize the low earth space debris environment down to sizes of 1 cm. The dual telescope staring survey, conducted in 1990, employed stereo (parallactic) viewing to enable estimates of the target altitudes to be determined. We will described the stereo survey techniques, a pipeline image processor to reduce the video tapes, and our plans for data analyses. We will also present a brief overview of the program results to date, emphasizing extensive observations of two small optically detected debris.

The Air Force Phillips Laboratory (PL) is tasked by Air Force Space Command to characterize the orbital debris environment. Part of this task is to search for and detect debris using the optical facilities at the PL Air Force Maui Optical Station (AMOS), which is located at the Maui Space Surveillance Site (MSSS). The goals of the program are discussed, with emphasis on the detection program. This includes telescopes and sensors available, how they are used, and handoffs from one sensor to another. Results of this correlation, as well as conclusions on the orbital debris environment, are presented.

The Air Force Phillips Laboratory is conducting measurements to characterize the orbital debris environment using wide-field optical systems located at the Air Force's Maui, Hawaii, Space Surveillance Site. Conversion of the observed visible brightnesses of detected debris objects to physical sizes require knowledge of the albedo (reflectivity). A thermal model for small debris objects has been developed and is used to calculate albedos from simultaneous visible and thermal infrared observations of catalogued debris objects. The model and initial results will be discussed.

Detection and measurement of small space debris objects are vital to verify the validity of debris models for the low Earth orbit (LEO) environment. Calibration of optical instruments is necessary so that reliable estimates of the size and albedo of man-made orbiting objects can be found. The Orbital Debris Radar Calibration Spheres (ODERACS) project is being conducted by NASA and the DoD to calibrate both radar and optical tracking facilities for small objects. This paper discusses the pre-flight optical calibration of the spheres. The purpose of this study is to determine the spectral reflectivity, scattering characteristics and albedo for the visible wavelength region. The measurements are performed by illuminating the flight spheres with a collimated beam of light, and measuring the reflected visible light over possible phase angles. This allows one to estimate the specular and scattering characteristics as well as the albedo. Tests were conducted on several flight and test metal spheres with varying diameters and surface characteristics. The polished metal spheres are shown to be very good specular reflectors, while the diffuse surfaces exhibit both specular and scattering reflection characteristics.

Laboratory investigations and development of space-based ultraviolet sensors have been done in Russia for over 20 years now. We have developed and continue to perfect the UV sensitive photocathodes, photosensitive transducers with luminescent screen for visualization of the UV image. Development of highly sensitive time-coordinate photoreceiving sensors is under way now including those on the basis of microchannel plates (MCP) operating in the photoelectrocounting mode, where photocathodes are employed. A peculiarity of this development is the fusion of photoreceiving unit with high performance computers, a new technology for registration of the enhanced photo current and new methods for computation and identification of the moving low contrast objects. Integral photoreceiving units with the very narrow selectivity by frequency have been developed on the basis of matrix chambers and photometers. For operation in the UV band CCD arrays are also used but their sensitivity is lower than that of the microchannel plates. It is possible to compensate the disadvantages of the CCD and the MCP devices by employment of photoelectron structures with memory, which are now at the stage of theoretic validation and laboratory research. Results achieved in development of key elements for new UV sensors and the experience of space studies by means of orbital sensors enable us to move to a new concept involving the use of small size spacecraft for the ecological monitoring of the earth and the near Earth space including their employment for surveillance of the small-size space debris.

We describe a sensitive technique for detecting small space debris that exploits a fast photon- counting imager. Microchannel plate detectors using crossed delay-line readout can achieve a resolution of 2048 X 2048 spatial pixels and a maximum count rate of about 10⁶ photons per second. A baseline debris-tracking system might couple this detector to a 16-cm aperture telescope. The detector yields x, y, and time information for each detected photon. When visualized in (x,y,t) space, photons from a fast-moving orbital object appear on a straight line. They can be distinguished from diffuse background photons, randomly scattered in the space, and star photons, which fall on a line with sidereal velocity. By searching for this unique signature, we can detect and track small debris objects. At dawn and dusk, a spherical object of 1.3 cm diameter at 400 km will reflect sunlight for an apparent magnitude of V approximately equals 16. The baseline system would detect about 16 photons from this object as it crosses a 1 degree field of view in about 1 second. The line in (x,y,t) space will be significant in a diffuse background of approximately 10⁶ photons. We discuss the data processing scheme and line detection algorithm. The advantages of this technique are that one can (1) detect cm-size debris objects with a small telescope, and (2) detect debris moving with any direction and velocity.

The results of 69 hours of zenith-staring search with the USAF GEODSS telescopes in 1990 is discussed. The flux rate is 7.2 objects/hour, roughly 10% less than that found from 1989 data. After correlation with objects in the Space Command Catalog, the ratio of cataloged objects to the total population is 0.56 versus a value of 0.45 found in 1989. Significant differences are seen in the data between the two observing sites, Maui and Diego Garcia, and between the three sets of objects resulting from the correlation with catalog data.

An accurate problem of the short and long term man-made orbital debris environment depends on the fidelity of spacecraft breakup models. These models provide post-breakup information for spacecraft collisions and explosions, including the number of fragments produced and the state vectors for each fragment. One way to verify the predictive capabilities of these models is to compare their predictions with the results of well characterized hypervelocity impact experiments. This paper compares the predictions of the empirically based spacecraft breakup model Impact 2.02 to the outcome of several hypervelocity impact experiments. The experimental data used in this analysis came from both ground- and space-based experiments. The experimental data and the model predictions are analyzed and correlated, and the discrepancies are identified.

This paper examines the vulnerability of the international Space Station Freedom (SSF) to impacts by orbital debris. Impact by debris particles with diameters of 1 cm and greater upon certain components of SSF would cause the failure of that component. NASA has calculated the frequency of 1 cm diameter and larger particle impacts upon SSF using its orbital debris flux model (found in NASA document SSP 30425). This impact frequency has been used to calculate component reliability based on the probability that a component will be impacted within the given station lifetime and suffer a catastrophic failure. This paper examined NASA's use of the orbital debris flux model in these calculations and proposes that conservative estimates within the orbital debris flux model may cause underestimation of SSF component reliabilities. Probability of collision (PC) calculations based upon recent U.S. Space Command satellite catalog information coupled with reasonable multiplication factors for 1 cm debris are presented, which yield lower impact frequencies than those found using the orbital debris flux model. The resulting lower probabilities of collision result in higher calculated values for component reliabilities; values that compare favorably with the proposed NASA safety criteria. Possible reasons for the disagreement between methods are discussed.

The J. N. Johnson void growth fracture model has been implemented into Magi to develop a capability to predict orbital debris generated from hypervelocity impacts or programmed explosions. Magi is a hydrodynamic code based on Smoothed Particle Hydrodynamics (SPH). This paper demonstrates advances in modeling fragmentation by calculating the debris environment generated from aluminum spheres impacting thin aluminum sheets representing debris shields. The shape of the debris cloud, fragment size, and fragment velocities are compared to the experimental work documented by Piekutowski. Plots of fragment mass versus cumulative number of fragments are also presented. Calculations of flyer plate experiments are presented and discussed to provide a foundation for understanding the fracture model and its input parameters. The results show that SPH is quite natural for modeling the fragmentation for these experiments.

Very high driving pressures (tens or hundreds of GPa), are required to accelerate flier plates to hypervelocities. This loading pressure pulse on the flier plates must be nearly shockless to prevent the plate from melting or vaporizing. This is accomplished by using graded-density impactors referred to as 'pillows'. When this graded-density material is used to impact a flier- plate in a modified two-stage light gas gun, nearly shockless megabar pressures are introduced into the flier plate. The pressure pulses must also be tailored to prevent spallation of the flier- plate. This technique has been used to launch nominally 1-mm-thick aluminum, magnesium and titanium (gram-size) intact plates to 10.4 km/s, and 0.5-mm-thick aluminum and titanium (half-gram size) intact plates to 12.2 km/s. This is the highest mass-velocity capability attained with laboratory launchers to date, and should open up new regimes of impact physics and lethality studies related to space sciences for laboratory investigations. In particular, the mass- velocity capability of this newly developed hypervelocity launcher meets the average specifications of the space debris environment, and is therefore expected to be a useful tool to evaluate the effects of debris impact on space structures and debris shields.

Well-controlled hypervelocity impact experiments, conducted in support of the Space Station Shielding Program, were numerically modeled using two sophisticated hydrocodes, the multi- dimensional hydrodynamics code CTH and the Smoothed Particle Hydrodynamics code (SPH). The experiments simulated the impact of space debris on single and double (Whipple shield) plate configurations. Impact velocities on the order of 10 km/s were applied to gram sized flier plates and spherical projectiles that struck thin (less than 1 cm) aluminum and steel plates. Computational predictions of the debris cloud dynamics and plate damage for these experiments were analyzed and correlated with the data obtained from pulsed laser photographs.

The orbital debris environment today is the consequence of explosion breakup events in both Low Earth Orbit and geosynchronous orbit. In the recent past, most operators have adopted practices which reduce the likelihood of further explosions. In the future, the debris environment is likely to be determined by collision events which become increasingly probable as the number of objects in orbit continues to grow. Limiting the likelihood of collisions by limiting the orbital lifetime of objects is more costly than the operational modifications that have been implemented to date, but appears to be necessary. Design and operations implications are examined.

Safety in space will become an increasingly important issue as the number of foreign space programs grow, and as the U.S. private sector increases its space activities. The U.S. Department of Transportation (DOT) is the federal agency responsible, under the Commercial Space launch Act of 1984, with regulation of the U.S. commercial space transportation industry in order to protect public health and safety, and safety of property. This paper discusses how the regulatory and licensing responsibilities of the Office of Commercial Space Transportation influence the safety of private sector launch operations in space. Of particular interest are impacts and benefits for commercial space services providers resulting from the government's safety research and technical safety evaluations.

The rapid increase in number and complexity of space tests in the future requires unique hazard analyses, test monitoring, and positive safety control. The Detachment 2 (Det 2) of the Space and Missiles System Control (SMC), formerly known as Consolidated Space Test Center (CSTC), at Onizuka AFB in Sunnyvale California provides an integrated space safety analysis and independent assessment capability required for certification of space safety test readiness and real-time control of hazardous space tests. The objective of this organization is to eliminate unnecessary risk to manned missions in space, national space assets, ground population and property, the space environment, and test articles.

Space debris pose a hazard to operations in outer space. This fact has been widely recognized by the scientific community. General measures necessary to minimize the hazard are well understood and have been clearly spelled out. The power to put relevant measures into force or to agree at least on unbinding recommendations rests, because of the global character of outer space, with international organizations, specifically with the United Nations. Up to 1992, proposals to put the question of space debris on the agenda of the UN Committee on the Peaceful Uses of Outer Space (COPUOS) have not been successful. In the meantime, the risk of fatal collisions of valuable satellites with space debris keeps increasing.

After nearly 35 years of space exploration and exploitation with little concern for the debris generated during routine spacecraft operations and upon the termination of space missions, concentrations of debris at certain orbital altitudes have reached alarming levels. Consequently, the risk of damage to operational spacecrafts in orbits with large debris populations can no longer be discounted. In the event the risk materializes and the source of debris can be identified, and if a cause of action exists, litigation could well ensue. By means of a hypothetical situation, this paper considers some of the possible causes of such litigation, as well as some of the legal issues the courts could be called upon to resolve.

Ten years ago, upper stages were left in space without any special precaution at the end of their mission. Due to propellant residuals on board, some of these stages (including an Ariane 1 third stage) exploded and produced an important amount of debris, sometimes in regions where the Space Debris Density was already critical. The problem has been solved for Ariane European launchers: general rules, aiming at launcher debris mitigation, are applied to Ariane 4 and 5 launchers design and operation. These cover from multiple launch capability to passivation of Ariane 4 and 5 upper stages. This policy, promoted by CNES in an international effort to establish rules on this subject, will likewise be continued for the Ariane 5 derivatives now under study.

On at least six occasions during 1983-1992, operational debris released from the fourth stage of Russian Proton launch vehicles fragmented, creating up to 60 new trackable debris in Earth orbit after each event. Surprisingly, these fragmentations occurred 18 - 96 months following successful Proton missions. One month after the fifth incident in September, 1992, an international investigation employing American space surveillance data and analyses and Russian engineering knowledge determined the probable cause of the satellite breakups. Preventive measures are now being developed for future Proton flights. The unprecedented Russian-American cooperation leading to the resolution of this environmental issue should serve as a model for future investigations.

Although the dangers from orbital debris are becoming more widely known, and orbital debris experts generally agree on the need for concerted international attention to the issue. Members of Congress need to be better informed about orbital debris concerns and how the nation might best resolve them. This paper examines the role that the U.S. Congress could play in the U.S. approach to reducing orbital debris. It also discusses the challenge posed by crafting an international solution and how Congress could assist in that important task.

Theoretical analyses of the orbital debris environment have shown that collisions could eventually become the dominant source of orbital debris. Debris will be added to the environment more rapidly by collisions than it can be removed by atmospheric drag. The models suggest that this process has already begun in some altitude regimes, specifically in the region from about 900 to 1000 km. It is important to verify these predictions by measurement in order to provide solid justification for early preventive measures. Collisions of two large objects are predicted to be very infrequent at the present time. However, the models predict that collisions of small debris with large objects will take place at an appreciable rate. One collision of an 0.5 cm object with a large object is expected each year. The collision rate increases with decreasing size, such that 80 collisions of 1 mm objects with a large object are expected each year. These collisions will not destroy the large object, but will generate a shower of microdebris. The particles generated by such collisions are mostly thrown forward, and either escape, or enter exceptionally long-lived elliptical orbits. It is suggested that an orbital experiment be flown for the specific purpose of detecting the predicted collisions by measuring the micro-debris fluxes resulting from the collisions.

Techniques for assessing the exposure of spacecraft surfaces to a limiting debris size and to penetration related characteristics such as crater diameter or surface penetration have been developed within a phase space definition of the debris environment. This approach requires the use of the full three dimensional description of the flux density distribution in the co- moving frame of the spacecraft. In this paper results of a simple modeling approach are used to generate debris environment and penetration related characteristics for the manmade debris environment as defined in the Space Station Natural and Induced Environment Document (SSP 30425), for a manmade environment derived of USSPACECOMMAND catalog element sets, and for the reference meteoroid environment in SSP 30425. Assuming the spacecraft is protected by an aluminum Whipple two-sheet shield, techniques for reducing shielding weight for a given reliability will be presented and results using this technique for several convex structural shape will be discussed.

Modeling of the space debris environment is critical to providing spacecraft designers with projections of the environment over the planned lifetime of a system. Understanding how the debris environment is projected to change over time allows spacecraft designers and managers the ability to implement responsive design mitigation measures into their systems. This paper develops a strategy for assessing the projected threat of the space debris environment for particular space systems. First, a historical simulation of the debris environment is made using available empirical models. For the system under study, the planned operational loss rates are also identified based on mission operational requirements and projected launch mission models. Generally, the system designers will plan to replace components at a matching rate to maintain required system performance. A threshold debris growth factor is then identified giving the percentage increase over the current historical environment that will produce a debris collision rate equal the planned spacecraft loss rate at identified from the mission operational requirements. Using this approach, the overall threat of the projected debris environment for particular space systems is couched in terms spacecraft managers can appreciate, that of replacements required over a given period of time.

NASA concern with the continuing growth in the orbital debris population has generated interest in developing a policy to limit orbital debris generation. To support this policy position a handbook is being prepared to enable developing NASA program to evaluate the effect of the debris environment on their programs and to enable them to assess the debris consequences of their program. In support of this handbook development, EVOLVE, a debris environment evolution code at NASA/JSC, was adapted to provide environment projections when debris mitigation measures are introduced. Results of this model development and application will be presented in this paper. Results from five future space usage scenarios will be discussed. One of the suggested mitigation measures for future space programs is that all hardware be removed from orbit within a short time after completion of mission. Such a procedure would have considerable impact on future space programs.

The IRIDIUM^TM Program represents the most exciting prospect in space operations in the world today. However, it also represents one of the biggest engineering challenges as well as one of the best commercial opportunities ever. This paper will describe the trend setting approach Motorola has taken in orbital debris mitigation. It will delineate the attributes of a comprehensive program being implemented during the design phase and to 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 mitigation.

This paper documents a parametric analysis of collision probability between any low Earth orbit space vehicle and orbital debris. Its purpose is to understand the probability of debris impact (risk) and assess how that risk can be managed through debris avoidance maneuvers by a space vehicle over a period of time. No matter how high the frequency of debris avoidance maneuvers, probability of debris impact (risk) can never be completely eliminated. The probability of collision over time, predicted state vector accuracies for debris and space vehicle, and maneuver rate for any space vehicle in low Earth orbit are fundamentally interrelated. This work discusses the rationale for the selection of the values of analysis parameters, the methods of determining hit probability for a close passage (conjunction) between a space vehicle and a debris object (conjunction), the method of determining maneuver rate and remaining risk, and the significance of the maneuver rate versus remaining risk curves. Assuming the Kessler orbital debris environment model1 and specific space vehicle dimensions, estimated remaining risk and maneuver rates are calculated.