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NASA NTRS2006

Earth Satellite Population Instability: Underscoring the Need for Debris Mitigation

A recent study by NASA indicates that the implementation of international orbital debris mitigation measures alone will not prevent a significant increase in the artificial Earth satellite population, beginning in the second half of this century. Whereas the focus of the aerospace community for the past 25 years has been on the curtailment of the generation of long-lived orbital debris, active remediation of the current orbital debris population should now be reconsidered to help preserve near-Earth space for future generations. In particular, we show in this paper that even if launch operations were to cease today, the population of space debris would continue to grow. Further, proposed remediation techniques do not appear to offer a viable solution. We therefore recommend that, while the aerospace community maintains the current debris-limiting mission regulations and postmission disposal procedures, future emphasis should be placed on finding new remediation technologies for solving this growing problem. Since the launch of Sputnik 1, space activities have created an orbital debris environment that poses increasing impact risks to existing space systems, including human space flight and robotic missions (1, 2). Currently, more than 9,000 Earth orbiting man-made objects (including many breakup fragments), with a combined mass exceeding 5 million kilograms, are tracked by the US Space Surveillance Network and maintained in the US satellite catalog (3-5). Three accidental collisions between cataloged satellites during the period from late 1991 to early 2005 have already been documented (6), although fortunately none resulted in the creation of large, trackable debris clouds. Several studies conducted during 1991-2001 demonstrated, with assumed future launch rates, the unintended growth potential of the Earth satellite population, resulting from random, accidental collisions among resident space objects (7-13). In some low Earth orbit (LEO) altitude regimes where the number density of satellites is above a critical spatial density, the production rate of new satellites (i.e., debris) due to collisions exceeds the loss of objects due to orbital decay. NASA s evolutionary satellite population model LEGEND (LEO-to-GEO Environment Debris model), developed by the Orbital Debris Program Office at the NASA Lyndon B. Johnson Space Center, is a high fidelity three-dimensional physical model that is capable of simulating the historical satellite environment, as well as the evolution of future debris populations (14, 15). The subject study assumed no rocket bodies and spacecraft were launched after December 2004, and no future disposal maneuvers were allowed for existing spacecraft, few of which currently have such a capability. The rate of satellite explosions would naturally decrease to zero within a few decades as the current satellite population ages. The LEGEND future projection adopts a Monte Carlo approach to simulate future on-orbit explosions and collisions. Within a given projection time step, once the explosion probability is estimated for an intact object, a random number is drawn and compared with the probability to determine if an explosion would occur. A similar procedure is applied to collisions for each pair of target and projectile involved within the same time step. Due to the nature of the Monte Carlo process, multiple projection runs must be performed and analyzed before one can draw reliable and meaningful conclusions from the outcome. A total of fifty, 200-year future projection Monte Carlo simulations were executed and evaluated (16).

Liou, Jer-chyi↗

NASA NTRSDate not supplied

Analysis of Hypervelocity-Impacted Thin Films for Space Applications

Exposure to the harsh space environment and hypervelocity impacts from micrometeoroids and small orbital debris can affect space operations through long-term degradation of spacecraft materials, surfaces, or systems. Multi-layer insulation (MLI) and coated polyimide films are space-craft materials commonly found on the external surfaces of spacecraft for thermal control and protection. Characterizing the damage to these exposed materials with in-situ and laboratory measurements can better inform spacecraft designers and operators in understanding and mitigating surface or system degradation. In this paper, we examine flown Hubble Space Telescope (HST) Electronic Bay 5 MLI and coated polyimide ground test articles to characterize hyper-velocity impact damage. Impact feature cavities are inspected to identify damage, which can validate long-term degradation models, improve thermal management systems, or improve accuracy of damage predictions. An overview of the impact characterization using optical microscopy, narrow-band spectroscopy, and scanning election microscopy/energy dispersive x-ray (SEM/EDX) of the two materials is presented. Observations of robustness in the space environment, as well as a detailed assessment of cratering and penetration statistics for the space-exposed sample, also are discussed.

Melissa A Ward↗

NASA NTRS2008

The Use of Solar Heating and Heat Cured Polymers for Lunar Surface Stabilization

Dust ejecta can affect visibility during a lunar landing, erode nearby coated surfaces and get into mechanical assemblies of in-place infrastructure. Regolith erosion was observed at many of the Apollo landing sites. This problem needs to be addressed at the beginning of the lunar base missions, as the amount of infrastructure susceptible to problems will increase with each landing. Protecting infrastructure from dust and debris is a crucial step in its long term functionality. A proposed way to mitigate these hazards is to build a lunar launch pad. Other areas of a lunar habitat will also need surface stabilization methods to help mitigate dust hazards. Roads would prevent dust from being lifted during movement and dust free zones might be required for certain areas critical to crew safety or to critical science missions. Work at NASA Kennedy Space Center (KSC) is investigating methods of stabilizing the lunar regolith including: sintering the regolith into a solid and using heat or UV cured polymers to stabilize the surface. Sintering, a method in which powders are heated until fusing into solids, has been proposed as one way of building a Lunar launch/landing pad. A solar concentrator has been built and used in the field to sinter JSC-1 Lunar stimulant. Polymer palliatives are used by the military to build helicopter landing pads and roads in dusty and sandy areas. Those polymers are dispersed in a solvent (water), making them unsuitable for lunar use. Commercially available, solvent free, polymer powders are being investigated to determine their viability to work in the same way as the solvent borne terrestrial analog. This presentation will describe the ongoing work at KSC in this field. Results from field testing will be presented. Physical testing results, including compression and abrasion, of field and laboratory prepared samples will be presented.

Hintze, Paul↗

NASA NTRS2019

NASA & US Government Orbital Debris Mitigation Policies

NASA's Orbital Debris Program Office (ODPO) is a global center of expertise in all aspects of orbital debris and spacecraft entry. The U.S. Government's regulations in such matters are subject to government-wide coordinated guidelines, but enforced at the agency level. Government-to-Government coordination of regulations and global agreements is through the IADC and UN. The U.S. Government policy for orbital debris has undergone a major update within the last two weeks to respond to the evolving space environment.

Bacon, Jack↗

NASA NTRS2014

NASA Tech Briefs, March 2014

Topics include: Data Fusion for Global Estimation of Forest Characteristics From Sparse Lidar Data; Debris and Ice Mapping Analysis Tool - Database; Data Acquisition and Processing Software - DAPS; Metal-Assisted Fabrication of Biodegradable Porous Silicon Nanostructures; Post-Growth, In Situ Adhesion of Carbon Nanotubes to a Substrate for Robust CNT Cathodes; Integrated PEMFC Flow Field Design for Gravity-Independent Passive Water Removal; Thermal Mechanical Preparation of Glass Spheres; Mechanistic-Based Multiaxial-Stochastic-Strength Model for Transversely-Isotropic Brittle Materials; Methods for Mitigating Space Radiation Effects, Fault Detection and Correction, and Processing Sensor Data; Compact Ka-Band Antenna Feed with Double Circularly Polarized Capability; Dual-Leadframe Transient Liquid Phase Bonded Power Semiconductor Module Assembly and Bonding Process; Quad First Stage Processor: A Four-Channel Digitizer and Digital Beam-Forming Processor; Protective Sleeve for a Pyrotechnic Reefing Line Cutter; Metabolic Heat Regenerated Temperature Swing Adsorption; CubeSat Deployable Log Periodic Dipole Array; Re-entry Vehicle Shape for Enhanced Performance; NanoRacks-Scale MEMS Gas Chromatograph System; Variable Camber Aerodynamic Control Surfaces and Active Wing Shaping Control; Spacecraft Line-of-Sight Stabilization Using LWIR Earth Signature; Technique for Finding Retro-Reflectors in Flash LIDAR Imagery; Novel Hemispherical Dynamic Camera for EVAs; 360 deg Visual Detection and Object Tracking on an Autonomous Surface Vehicle; Simulation of Charge Carrier Mobility in Conducting Polymers; Observational Data Formatter Using CMOR for CMIP5; Propellant Loading Physics Model for Fault Detection Isolation and Recovery; Probabilistic Guidance for Swarms of Autonomous Agents; Reducing Drift in Stereo Visual Odometry; Future Air-Traffic Management Concepts Evaluation Tool; Examination and A Priori Analysis of a Direct Numerical Simulation Database for High-Pressure Turbulent Flows; and Resource-Constrained Application of Support Vector Machines to Imagery.

Source record↗

NASA NTRS2013

Independent Assessment of Instrumentation for ISS On-Orbit NDE

International Space Station (ISS) Structural and Mechanical Systems Manager, requested that the NASA Engineering and Safety Center (NESC) provide a quantitative assessment of commercially available nondestructive evaluation (NDE) instruments for potential application to the ISS. This work supports risk mitigation as outlined in the ISS Integrated Risk Management Application (IRMA) Watch Item #4669, which addresses the requirement for structural integrity after an ISS pressure wall leak in the event of a penetration due to micrometeoroid or debris (MMOD) impact. This document contains the outcome of the NESC assessment.

Madaras, Eric I↗

NASA NTRS2010

An Assessment of the Current LEO Debris Environment and the Need for Active Debris Removal

The anti-satellite test on the Fengun-1 C weather satellite in early 2007 and the collision between Iridium 33 and Cosmos 2251 in 2009 dramatically altered the landscape of the human-made orbital debris environment in the low Earth orbit (LEO). The two events generated approximately 5500 fragments large enough to be tracked by the U.S. Space Surveillance Network. Those fragments account for more than 60% increase to the debris population in LEO. However, even before the ASAT test, model analyses already indicated that the debris population (for those larger than 10 cm) in LEO had reached a point where the population would continue to increase, due to collisions among existing objects, even without any future launches. The conclusion implies that as satellites continue to be launched and unexpected breakup events continue to occur, commonly-adopted mitigation measures will not be able to stop the collision-driven population growth. To remediate the debris environment in LEO, active debris removal must be considered. This presentation will provide an updated assessment of the debris environment after the Iridium 33/Cosmos 2251 collision, an analysis of several future environment projections based on different scenarios, and a projection of collision activities in LEO in the near future. The need to use active debris removal to stabilize future debris environment will be demonstrated and the effectiveness of various active debris removal strategies will be quantified.

Liou, Jer-Chyi↗

NASA NTRS2010

Experimental Space Shuttle Orbiter Studies to Acquire Data for Code and Flight Heating Model Validation

In an experimental study to obtain detailed heating data over the Space Shuttle Orbiter, CUBRC has completed an extensive matrix of experiments using three distinct models and two unique hypervelocity wind tunnel facilities. This detailed data will be employed to assess heating augmentation due to boundary layer transition on the Orbiter wing leading edge and wind side acreage with comparisons to computational methods and flight data obtained during the Orbiter Entry Boundary Layer Flight Experiment and HYTHIRM during STS-119 reentry. These comparisons will facilitate critical updates to be made to the engineering tools employed to make assessments about natural and tripped boundary layer transition during Orbiter reentry. To achieve the goals of this study data was obtained over a range of Mach numbers from 10 to 18, with flight scaled Reynolds numbers and model attitudes representing key points on the Orbiter reentry trajectory. The first of these studies were performed as an integral part of Return to Flight activities following the accident that occurred during the reentry of the Space Shuttle Columbia (STS-107) in February of 2003. This accident was caused by debris, which originated from the foam covering the external tank bipod fitting ramps, striking and damaging critical wing leading edge heating tiles that reside in the Orbiter bow shock/wing interaction region. During investigation of the accident aeroheating team members discovered that only a limited amount of experimental wing leading edge data existed in this critical peak heating area and a need arose to acquire a detailed dataset of heating in this region. This new dataset was acquired in three phases consisting of a risk mitigation phase employing a 1.8% scale Orbiter model with special temperature sensitive paint covering the wing leading edge, a 0.9% scale Orbiter model with high resolution thin-film instrumentation in the span direction, and the primary 1.8% scale Orbiter model with detailed thin-film resolution in both the span and chord direction in the area of peak heating. Additional objectives of this first study included: obtaining natural or tripped turbulent wing leading edge heating levels, assessing the effectiveness of protuberances and cavities placed at specified locations on the orbiter over a range of Mach numbers and Reynolds numbers to evaluate and compare to existing engineering and computational tools, obtaining cavity floor heating to aid in the verification of cavity heating correlations, acquiring control surface deflection heating data on both the main body flap and elevons, and obtain high speed schlieren videos of the interaction of the orbiter nose bow shock with the wing leading edge. To support these objectives, the stainless steel 1.8% scale orbiter model in addition to the sensors on the wing leading edge was instrumented down the windward centerline, over the wing acreage on the port side, and painted with temperature sensitive paint on the starboard side wing acreage. In all, the stainless steel 1.8% scale Orbiter model was instrumented with over three-hundred highly sensitive thin-film heating sensors, two-hundred of which were located in the wing leading edge shock interaction region. Further experimental studies will also be performed following the successful acquisition of flight data during the Orbiter Entry Boundary Layer Flight Experiment and HYTHIRM on STS-119 at specific data points simulating flight conditions and geometries. Additional instrumentation and a protuberance matching the layout present during the STS-119 boundary layer transition flight experiment were added with testing performed at Mach number and Reynolds number conditions simulating conditions experienced in flight. In addition to the experimental studies, CUBRC also performed a large amount of CFD analysis to confirm and validate not only the tunnel freestream conditions, but also 3D flows over the orbiter acreage, wing leading edge, and controlurfaces to assess data quality, shock interaction locations, and control surface separation regions. This analysis is a standard part of any experimental program at CUBRC, and this information was of key importance for post-test data quality analysis and understanding particular phenomena seen in the data. All work during this effort was sponsored and paid for by the NASA Space Shuttle Program Office at the Johnson Space Center in Houston, Texas.

Wadhams, T. P.↗

NASA NTRS2026

Space Mission Options for Reconnaissance and Mitigation of Asteroid 2024 YR4

Near-Earth asteroid 2024 YR 4 was discovered on 2024-12-27 and its probability of Earth impact in December 2032 peaked at about 3% on 2025-02-18. Additional observations ruled out Earth impact by 2025-02-23. However, the probability of lunar impact in December 2032 then rose, reaching about 4% by the end of the apparition in May 2025. James Webb Space Telescope (JWST) observations on 2025-03-26 estimated the asteroid's diameter at 60 +/- 7 m. Studies of 2024 YR 4 's potential lunar impact effects suggest lunar ejecta could increase micrometeoroid debris flux in low Earth orbit up to 1000 times above background levels over just a few days, possibly threatening astronauts and spacecraft. In this work, we present options for space missions to 2024 YR 4 that could be utilized if lunar impact is confirmed. We cover flyby & rendezvous reconnaissance, deflection, and robust disruption of the asteroid. We examine both rapid-response and delayed launch options through 2032. We evaluate chemical and solar electric propulsion, various launch vehicles, optimized deep space maneuvers, and gravity assists. Re-tasking extant spacecraft and using built spacecraft not yet launched are also considered. The best reconnaissance mission options launch in late 2028, leaving only approximately three years for development at the time of this writing in August 2025. Deflection missions were assessed and appear impractical. However, kinetic robust disruption missions are available with launches between April 2030 and April 2032. Nuclear robust disruption missions are also available with launches between late 2029 and late 2031. Finally, even if lunar impact is ruled out there is significant potential utility in deploying a reconnaissance mission to characterize the asteroid.

Asteroid Disruption↗

NASA NTRS2020

Recent Radar Observations of the Sub-Centimeter Orbital Debris Environment

The NASA Orbital Debris Program Office (ODPO) has conducted radar observations of the orbital debris environment since the early 1990’s to provide measurement data that supports orbital debris models and risk mitigation activities in support of NASA mission objectives. Orbital debris radar observations are a unique mode for radar operation, employing a fixed beam configuration to statistically sample the environment. An advantage of conducting operations in this fashion is that it enables observations of smaller classes of orbital debris than would otherwise be available from the same sensor operating in a traditional tracking mode. Orbital debris-mode radar observations are used to fill in the gaps, which exist in the currently available data from the Space Surveillance Network (SSN), on small size orbital debris populations that represent significant risk to NASA programs. These gaps have typically covered orbital debris with characteristic sizes less than approximately 10 cm down to approximately 3 mm in low Earth orbit (LEO) – depending upon the altitude and sensor configuration. The value of orbital debris radar measurements lies in the ability to extract partial orbital element information about orbital debris in the centimeter to several millimeter size regimes in low Earth orbit – which are not available from other measurement sources. This paper will discuss observations of this smaller class of orbital debris observed in recent years from the radars at the MIT Haystack Observatory in Westford, Massachusetts, and the Goldstone Solar System Radar near Barstow, California. The former radar is able to observe orbital debris down to approximately 5 mm, and the latter, orbital debris with characteristic sizes near 3 mm – at altitudes less than 1000 km. The characteristics and inferences about the current LEO orbital debris environment, and the different subpopulations that are identifiable in the observations are highlighted.

Kennedy, Timothy↗

NASA NTRS1990

Space environments and their effects on space automation and robotics

Automated and robotic systems will be exposed to a variety of environmental anomalies as a result of adverse interactions with the space environment. As an example, the coupling of electrical transients into control systems, due to EMI from plasma interactions and solar array arcing, may cause spurious commands that could be difficult to detect and correct in time to prevent damage during critical operations. Spacecraft glow and space debris could introduce false imaging information into optical sensor systems. The presentation provides a brief overview of the primary environments (plasma, neutral atmosphere, magnetic and electric fields, and solid particulates) that cause such adverse interactions. The descriptions, while brief, are intended to provide a basis for the other papers presented at this conference which detail the key interactions with automated and robotic systems. Given the growing complexity and sensitivity of automated and robotic space systems, an understanding of adverse space environments will be crucial to mitigating their effects.

Garrett, Henry B.↗

NASA NTRS2007

Space Shuttle Propulsion Systems Plume Modeling and Simulation for the Lift-Off Computational Fluid Dynamics Model

This paper details advances being made in the development of Reynolds-Averaged Navier-Stokes numerical simulation tools, models, and methods for the integrated Space Shuttle Vehicle at launch. The conceptual model and modeling approach described includes the development of multiple computational models to appropriately analyze the potential debris transport for critical debris sources at Lift-Off. The conceptual model described herein involves the integration of propulsion analysis for the nozzle/plume flow with the overall 3D vehicle flowfield at Lift-Off. Debris Transport Analyses are being performed using the Shuttle Lift-Off models to assess the risk to the vehicle from Lift-Off debris and appropriately prioritized mitigation of potential debris sources to continue to reduce vehicle risk. These integrated simulations are being used to evaluate plume-induced debris environments where the multi-plume interactions with the launch facility can potentially accelerate debris particles toward the vehicle.

Strutzenberg, L. L.↗

NASA NTRS2010

The Successful Implementation of NASA Orbital Debris Requirements for the Retirement of TDRS-1

TDRS-1 was decommissioned on October 28th 2009 following more than 26 years of operation. The Grand Old Dame of the TDRSS constellation wa s launched aboard the maiden voyage of the Space Shuttle Challenger ( STS-6) in April 1983. TDRS-1 survived a malfunction of the Inertial Upper Stage eventually utilizing its own propulsion system to success fully reach its assigned station in geosynchronous orbit. The anomalo us beginning of the TDRS-1 mission was not without lasting consequenc es as the primary reaction control system (A-side) was completely di sabled with an apparent propellant leak and the secondary system (B-s ide) suffered damage to its negative roll thruster rendering the thru ster inoperable. Following decommissioning the challenge to completin g a successful TDRS-1 end-of-mission (EOM) was the implementation of the stringent orbital debris requirements of NPR 8715.6 with a parti ally functioning spacecraft not originally designed to meet those req uirements. The TDRS-1 EOM had three key goals: 1) removal of the spac ecraft from geosynchronous orbit; 2) depletion of the remaining prope llant; and 3) passivation of all other sources of energy storage or generation. The TDRS-1 EOM approach was one of minimizing risks while accomplishing the goals above. The orbit raising portion of EOM was accomplished using deltavelocity operations already proven during pre vious stationchanging maneuvers. The propellant depletion approach wa s necessarily more aggressive as over 20 hours of burn time was requ ired to deplete the remaining fuel. A novel approach utilizing a spin ning, thrusting, passively controlled spacecraft configuration was ut ilized to achieve reasonable burn durations that met schedule constra ints. This nonstandard configuration required careful analysis of ele ctrical, thermal, and communication subsystems. The configuration wa s thoroughly simulated prior to the start of operations and carefully characterized during the initial spin period and first burn. Passiva tion was by definition a unique operation not previously performed wi thin the TDRS 1-7 constellation. Use was made of a TDRS spacecraft si mulator to verify the operational procedure to mitigate risks and pro vide crew training. TDRS-1 orbit raising maneuvers commenced on June 5th 2010 and completed on June 14th with an apogee 370 km and a peri gee 352 km above geosynchronous altitude. 127 kg of propellant were e stimated to be remaining in the tanks at the completion of orbit rais ing. TDRS-1 was placed in its spinning orientation on June 16th and the first fuel depletion burn was performed the following day. A seri es of 10 depletion burns were performed ending on June 26th when both propellant tank pressures experienced dramatic drops. Final passivat ion was performed on June 27 th deactivating electronics, removing t he batteries from the bus and solar arrays, and disabling the space-t o-ground communications equipment. The Second TDRS Ground Terminal (S TGT) continued to open loop track TDRS-1 for several days attempting command reacquisition several times a day. All attempts were unsucce ssful confirming passivation was achieved. The TDRS-1 orbit at passiv ation was an orbital debris compliant 36,319 x 36,128 km in height. W hile differing spacecraft designs may preclude mimicking the exact TD RS-1 EOM approach, the TDRS-1 campaign serves to demonstrate that pr e-NPR 8715.6 designs can be made to meet the requirements resulting i n a reduced orbit debris environment for future missions

Mirczak, Walter↗

NASA NTRSDate not supplied

Design for Minimum Casualty Area – The IXPE Case

The Imaging X-Ray Polarimetry Explorer (IXPE) is a new international space observatory in NASA’s Small Explorer program, designed in a collaboration between the Italian Space Agency and NASA’s Marshall Space Flight Center, and built by Ball Aerospace. IXPE has an expected launch in November 2021, to a 600-km altitude equatorial orbit. IXPE is an astrophysics mission using three telescope assemblies to measure the polarization of cosmic X-rays. Each assembly is composed of a mirror module assembly (MMA) with 24 nested nickel-cobalt cylindrical shells and a unique, polarization-sensitive, gas pixel detector (GPD) within the detector unit (DU). As a NASA mission, IXPE must adhere to the orbital debris mitigation requirements specified in NASA Standard 8719.14 [1]; in the present work, we will only discuss reentry human casualty risk. As initially designed, the IXPE observatory exceeded NASA’s casualty risk threshold. IXPE does not include a propulsion system to perform a controlled reentry at the end of mission to mitigate the ground casualty risk. To reduce the risk from the uncontrolled reentry of this observatory, the IXPE design team worked with the NASA Orbital Debris Program Office to reduce the debris casualty area through design-for-demise and containment methods. The flight design of IXPE is now compliant with the ground casualty risk requirement at a casualty probability of 1:13,100 compared to a casualty requirement of better than 1:10,000.

C Ostrom↗

NASA NTRS2005

Range Systems Simulation for the NASA Shuttle: Emphasis on Disaster and Prevention Management During Lift-Off

This article describes a decision-making system composed of a number of safety and environmental models for the launch phase of a NASA Space Shuttle mission. The components of this distributed simulation environment represent the different systems that must collaborate to establish the Expectation of Casualties (E(sub c)) caused by a failed Space Shuttle launch and subsequent explosion (accidental or instructed) of the spacecraft shortly after liftoff. This decision-making tool employs Space Shuttle reliability models, trajectory models, a blast model, weather dissemination systems, population models, amount and type of toxicants, gas dispersion models, human response functions to toxicants, and a geographical information system. Since one of the important features of this proposed simulation environment is to measure blast, toxic, and debris effects, the clear benefits is that it can help safety managers not only estimate the population at risk, but also to help plan evacuations, make sheltering decisions, establish the resources required to provide aid and comfort, and mitigate damages in case of a disaster.

Rabelo, Lisa↗

NASA NTRS1995

Long term microparticle impact fluxes on LDEF determined from optical survey of Interplanetary Dust Experiment (IDE) sensors

Many of the IDE metal-oxide-silicon (MOS) capacitor-discharge impact sensors remained active during the entire Long Duration Exposure Facility (LDEF) mission. An optical survey of impact sites on the active surfaces of these sensors has been extended to include all sensors from the low-flux sides of LDEF (i.e. the west or trailing side, the earth end, and the space end) and 5-7 active sensors from each LDEF's high-flux sides (i.e. the east or leading side, the south side, and the north side). This survey was facilitated by the presence of a relatively large (greater than 50 micron diameter) optical signature associated with each impact site on the active sensor surfaces. Of the approximately 4700 impacts in the optical survey data set, 84% were from particles in the 0.5 to 3 micron size range. An estimate of the total number of hypervelocity impacts on LDEF from particles greater than 0.5 micron diameter yields a value of approximately 7 x 10(exp 6). Impact feature dimensions for several dozen large craters on MOS sensors and germanium witness plates are also presented. Impact fluxes calculated from the IDE survey data closely matched surveys of similar size impacts (greater than or equal to 3 micron diameter craters in Al, or marginal penetrations of a 2.4 micron thick Al foil) by other LDEF investigators. Since the first year IDE data were electronically recorded, the flux data could be divided into three long term time periods: the first year, the entire 5.8 year mission, and the intervening 4.8 years (by difference). The IDE data show that there was an order of magnitude decrease in the long term microparticle impact flux on the trailing side of LDEF, from 1.01 to 0.098 x 10(exp -4) m(exp 2)/s, from the first year in orbit compared to years 2-6. The long term flux on the leading edge showed an increase from 8.6 to 11.2 x 10(exp -4) m(exp -2)/s over this same time period. (Short term flux increases up to 10,000 times the background rate were recorded on the leading side during LDEF's first year in orbit.) The overall east/west ratio was 44, but during LDEF's first year in orbit the ratio was 8.5, and during years 2-6 the ratio was 114. Long term microparticle impact fluxes on the space end decreased from 1.12 to 0.55 x 10(exp -4) m(exp -2)/s from the first year in orbit compared to years 2-6. The earth end showed the opposite trend with an increase from 0.16 to 0.38 x 10(exp -4) m(exp -2)/s. Fluxes on rows 6 and 12 decreased from 6.1 to 3.4 and 6.7 to 3.7 x 10(exp -4) m(exp -2)/s, respectively, over the same time periods. This resulted in space/earth microparticle impact flux ratios of 7.1 during the first year and 1.5 during years 2-6, while the south/north, space/north and space/south ratios remained constant at 1.1, 0.16 and 0.17, respectively, during the entire mission. This information indicates the possible identification of long term changes in discrete microparticle orbital debris component contributions to the total impact flux experienced by LDEF. A dramatic decrease in the debris population capable of striking the trailing side was detected that could possibly be attributed to the hiatus of western launch activity experienced from 1986-1989. A significant increase in the debris population that preferentially struck the leading side was also observed and could possibly be attributed to a single breakup event that occurred in September of 1986. A substantial increase in the microparticle debris population that struck the earth end of LDEF, but not the space end, was also detected and could possibly be the result of a single breakup event at low altitude. These results point to the importance of including discrete orbital debris component contribution changes in flux models in order to achieve accurate predictions of the microparticle environment that a particular spacecraft will experience in earth orbit. The only reliable, verified empirical measurements of these changes are reported in this paper. Further time-resolved in-situ measurements of these debris populations are needed to accurately assess model predictions and mitigation practices.

Simon, C. G.↗

NASA NTRS1999

NASA Lewis Launch Collision Probability Model Developed and Analyzed

There are nearly 10,000 tracked objects orbiting the earth. These objects encompass manned objects, active and decommissioned satellites, spent rocket bodies, and debris. They range from a few centimeters across to the size of the MIR space station. Anytime a new satellite is launched, the launch vehicle with its payload attached passes through an area of space in which these objects orbit. Although the population density of these objects is low, there always is a small but finite probability of collision between the launch vehicle and one or more of these space objects. Even though the probability of collision is very low, for some payloads even this small risk is unacceptable. To mitigate the small risk of collision associated with launching at an arbitrary time within the daily launch window, NASA performs a prelaunch mission assurance Collision Avoidance Analysis (or COLA). For the COLA of the Cassini spacecraft, the NASA Lewis Research Center conducted an in-house development and analysis of a model for launch collision probability. The model allows a minimum clearance criteria to be used with the COLA analysis to ensure an acceptably low probability of collision. If, for any given liftoff time, the nominal launch vehicle trajectory would pass a space object with less than the minimum required clearance, launch would not be attempted at that time. The model assumes that the nominal positions of the orbiting objects and of the launch vehicle can be predicted as a function of time, and therefore, that any tracked object that comes within close proximity of the launch vehicle can be identified. For any such pair, these nominal positions can be used to calculate a nominal miss distance. The actual miss distances may differ substantially from the nominal miss distance, due, in part, to the statistical uncertainty of the knowledge of the objects positions. The model further assumes that these position uncertainties can be described with position covariance matrices. With these, and some additional simplifying assumptions, a closed-form solution for the probability of collision is obtained. This solution provides clear insights into how each of the independent parameters affects the probability of collision. It shows that for a given maximum probability of collision and prior knowledge of the objects position uncertainties and sizes, only knowledge of the nominal closest approach distance is required to make the launch/no launch decision. The model was completed and used for the mission assurance COLA analysis for the Cassini spacecraft, which was launched on a Titan IV/Centaur rocket on October 15, 1997. Although the model was specifically developed for the Cassini mission, it is clearly applicable for other launches as well. The effect of COLA closures on the launch window is shown. The bar represents the entire 140-min launch window on October 15, 1997; the blackened areas represent the loss of launch opportunities due to the potential for a collision with an orbiting object.

Bollenbacher, Gary↗

NASA NTRS2014

The Orbital Debris Problem and the Challenges for Environment Remediation

LEO debris population will continue to increase even with a good implementation of the commonly-adopted mitigation measures. The root-cause of the increase is catastrophic collisions involving large/massive intact objects (rocket bodies or spacecraft). The major mission-ending risks for most operational spacecraft, however, come from impacts with debris just above the threshold of the protection shields (~5-mm to 1-cm). A solution-driven approach is to seek: Concepts for removal of massive intacts with high P(collision); Concepts capable of preventing collisions involving intacts; Concepts for removal of 5-mm to 1-cm debris; Enhanced impact protection shields for valuable space assets. Key questions for remediation consideration of orbital debris: What is the acceptable threat level? What are the mission objectives? What is the appropriate roadmap/timeframe for remediation? Support advanced technology development when an economically viable approach is identified. Address non-technical issues, such as policy, coordination, ownership, legal, and liability at the national and international levels.

Liou, J.-C.↗