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All sourcesNASA NTRSDOE OSTI
NASA NTRS1991

Increased ephemeris accuracy using attitude-dependent aerodynamic force coefficients for inertially stabilized spacecraft

The FREEMAC program used to generate the aerodynamic coefficients, as well as associated routines that allow the results to be used in other software is described. These capabilities are applied in two numerical examples to the short-term orbit prediction of the Gamma Ray Observatory (GRO) and Hubble Space Telescope (HST) spacecraft. Predictions using attitude-dependent aerodynamic coefficients were made on a modified version of the PC-based Ephemeris Generation Program (EPHGEN) and were compared to definitive orbit solutions obtained from actual tracking data. The numerical results show improvement in the predicted semi-major axis and along-track positions that would seem to be worth the added computational effort. Finally, other orbit and attitude analysis applications are noted that could profit from using FREEMAC-calculated aerodynamic coefficients, including orbital lifetime studies, orbit determination methods, attitude dynamics simulators, and spacecraft control system component sizing.

Folta, David C.↗
NASA NTRS2009

The International Space Station and the Space Debris Environment: 10 Years On

For just over a decade the International Space Station (ISS), the most heavily protected vehicle in Earth orbit, has weathered the space debris environment well. Numerous hypervelocity impact features on the surface of ISS caused by small orbital debris and meteoroids have been observed. In addition to typical impacts seen on the large solar arrays, craters have been discovered on windows, hand rails, thermal blankets, radiators, and even a visiting logistics module. None of these impacts have resulted in any degradation of the operation or mission of the ISS. Validating the rate of small particle impacts on the ISS as predicted by space debris environment models is extremely complex. First, the ISS has been an evolving structure, from its original 20 metric tons to nearly 300 metric tons (excluding logistics vehicles) ten years later. Hence, the anticipated space debris impact rate has grown with the increasing size of ISS. Secondly, a comprehensive visual or photographic examination of the complete exterior of ISS has never been accomplished. In fact, most impact features have been discovered serendipitously. Further complications include the estimation of the size of an impacting particle without knowing its mass, velocity, and angle of impact and the effect of shadowing by some ISS components. Inadvertently and deliberately, the ISS has also been the source of space debris. The U.S. Space Surveillance Network officially cataloged 65 debris from ISS from November 1998 to November 2008: from lost cameras, sockets, and tool bags to intentionally discarded equipment and an old space suit. Fortunately, the majority of these objects fall back to Earth quickly with an average orbital lifetime of less than two months and a maximum orbital lifetime of a little more than 15 months. The cumulative total number of debris object-years is almost exactly 10, the equivalent of one piece of debris remaining in orbit for 10 years. An unknown number of debris too small to be tracked and cataloged have also been generated, but normally with even shorter orbital lifetimes. Finally, eight collision avoidance maneuvers have been performed to avoid potential collisions between ISS and large, tracked space debris. The most recent such maneuver was accomplished by ESA's Automated Transfer Vehicle, the Jules Verne, just three months before the 10th anniversary of the launch of ISS's first element.

Johnson, Nicholas↗
NASA NTRS2009

Launch Order, Launch Separation, and Loiter in the Constellation 1 1/2-Launch Solution

The NASA Constellation Program (CxP) is developing a two-element Earth-to-Orbit launch system to enable human exploration of the Moon. The first element, Ares I, is a human-rated system that consists of a first stage based on the Space Shuttle Program's solid rocket booster (SRB) and an upper stage that consists of a four-crew Orion capsule, a service module, and a Launch Escape System. The second element, Ares V, is a Saturn V-plus category launch system that consists of the core stage with a cluster of six RS-68B engines and augmented with two 5.5-segment SRBs, a Saturn-derived J-2X engine powering an Earth Departure Stage (EDS), and the lunar-lander vehicle payload, Altair. Initial plans called for the Ares V to be launched first, followed the next day by the Ares I. After the EDS performs the final portion of ascent and subsequent orbit circularization, the Orion spacecraft then performs a rendezvous and docks with the EDS and its Altair payload. Following checkout, the integrated stack loiters in low Earth orbit (LEO) until the appropriate Trans-Lunar Injection (TLI) window opportunity opens, at which time the EDS propels the integrated Orion Altair to the Moon. Successful completion of this 1 1/2-launch solution carries risks related to both the orbital lifetime of the assets and the probability of achieving the launch of the second vehicle within the orbital lifetime of the first. These risks, which are significant in terms of overall system design choices and probability of mission success, dictated a thorough reevaluation of the launch strategy, including the order of vehicle launch and the planned time period between launches. The goal of the effort described in this paper was to select a launch strategy that would result in the greatest possible expected system performance, while accounting for launch risks and the cost of increased orbital lifetime. Discrete Event Simulation (DES) model of the launch strategies was created to determine the probability of a second launch not occurring in a timely fashion (i.e., before the assets waiting in LEO expire). An overview of the launch strategy evaluation process is presented, along with results of specific cases that were analyzed. A high-level comparison of options is then presented, along with the conclusion derived from the analysis.

Stromgren, Chel↗
NASA NTRSDate not supplied

TechEdSat 7, 10, 13, 15: Exo-brake Experiments on the ISS, First Virgin Orbit, and First Firefly-Alpha Test Flights.

The TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has been studying cube satellite re-entry technologies with a focus on Exo-Brake drag device research. An exo-atmospheric braking device, the Exo-Brake uses the tenuous upper atmosphere acting in a free molecular flow regime to enable controllable adjustment of the drag profile of the host spacecraft, currently enabling rapid disposal of small spacecraft after mission conclusion, and eventual targeted de-orbit to a desired entry point at the Von Karman line, or approximately 100km in altitude. The TES-7 mission, flown on the first successful Virgin Orbit flight January 17, 2021, was injected into a 500km, 61-degree inclination orbit, with an expected orbital lifetime of approximately ten years. After successful deployment of a unique hydrogen gas-cell inflation design ‘disposal’-type Mylar Exo-Brake intended to rapidly de-orbit the spacecraft, the maximum expected orbital lifetime of the TES-7 spacecraft was successfully reduced from ten years to 1.3 years, an 87% reduction in orbit lifetime via a passive, fuel-less system. The subsequent TES-13 mission utilized a new, less complex ‘disposal’-type Exo-Brake design made from collapsible rigid struts rather than inflated struts to reduce the engineering, construction, and safety complexities introduced by the prior hydrogen gas inflation design. The rigid strut Exo-Brake design is stowed via rotational compression and deploys using only stored spring energy once released via electronic actuator. This new rotation compression storage design has become the new Exo-Brake design standard used on TES missions thanks to its simplicity. Following on the success of this new Exo-Brake design demonstrated by TES-13, the design was further refined to enabled active manipulation of the Exo-Brake effective drag area though a winch-like system. Both the TES-10 and TES-15 spacecraft were equipped with geometrically similar Exo-Brake devices capable of such active drag modulation, with the intent of studying the operational impact of having such devices on cube satellite missions. To increase the survival time of the Exo-Brake at low altitudes and thus gain more experiment time and better guidance capability, TES-15 was equipped with a modulated Exo-Brake constructed from 3M™ Nextel™ 440 ceramic oxide fiber, the same fabric material used in the Space Shuttle TPS system, rather than Mylar. After launching on Alpha Flight 2 October 1st, 2022, Firefly’s first successful launch, TES-15 met the same modulation restrictions as TES-10, with modulation prohibited above 200km to avoid possible collision with other spacecraft. In this case, a low deployment orbit caused an extremely short mission life, which, coupled with delayed spacecraft identification, exacerbated collision concerns and caused performance data of the high-temperature Exo-Brake design to be inconclusive. As such, the Nano Orbital Workshop has taken a new approach of only conducting modulated Exo-Brake research on missions deploying below ISS altitude to avoid collision avoidance restrictions, and is working to improve spacecraft identification and location reporting techniques to increase experiment durations. Upcoming TES mission will therefore use one of two classes of Exo-Brake depending on their target altitude, as to be described in the submission.

M Murbach↗
NASA NTRSDate not supplied

TechEdSat 7, 10, 13, 15: Exo-brake Experiments on the ISS, First Virgin Orbit, and First Firefly-Alpha Test Flights

The TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has been studying cube satellite re-entry technologies with a focus on Exo-Brake drag device research. An exo-atmospheric braking device, the Exo-Brake uses the tenuous upper atmosphere acting in a free molecular flow regime to enable controllable adjustment of the drag profile of the host spacecraft, currently enabling rapid disposal of small spacecraft after mission conclusion, and eventual targeted de-orbit to a desired entry point at the Von Karman line, or approximately 100km in altitude. The TES-7 mission, flown on the first successful Virgin Orbit flight January 17, 2021, was injected into a 500km, 61-degree inclination orbit, with an expected orbital lifetime of approximately ten years. After successful deployment of a unique hydrogen gas-cell inflation design ‘disposal’-type Mylar Exo-Brake intended to rapidly de-orbit the spacecraft, the maximum expected orbital lifetime of the TES-7 spacecraft was successfully reduced from ten years to 1.3 years, an 87% reduction in orbit lifetime via a passive, fuel-less system. The subsequent TES-13 mission utilized a new, less complex ‘disposal’-type Exo-Brake design made from collapsible rigid struts rather than inflated struts to reduce the engineering, construction, and safety complexities introduced by the prior hydrogen gas inflation design. The rigid strut Exo-Brake design is stowed via rotational compression and deploys using only stored spring energy once released via electronic actuator. This new rotation compression storage design has become the new Exo-Brake design standard used on TES missions thanks to its simplicity. Following on the success of this new Exo-Brake design demonstrated by TES-13, the design was further refined to enabled active manipulation of the Exo-Brake effective drag area though a winch-like system. Both the TES-10 and TES-15 spacecraft were equipped with geometrically similar Exo-Brake devices capable of such active drag modulation, with the intent of studying the operational impact of having such devices on cube satellite missions. To increase the survival time of the Exo-Brake at low altitudes and thus gain more experiment time and better guidance capability, TES-15 was equipped with a modulated Exo-Brake constructed from 3M™ Nextel™ 440 ceramic oxide fiber, the same fabric material used in the Space Shuttle TPS system, rather than Mylar. After launching on Alpha Flight 2 October 1st, 2022, Firefly’s first successful launch, TES-15 met the same modulation restrictions as TES-10, with modulation prohibited above 200km to avoid possible collision with other spacecraft. In this case, a low deployment orbit caused an extremely short mission life, which, coupled with delayed spacecraft identification, exacerbated collision concerns and caused performance data of the high-temperature Exo-Brake design to be inconclusive. As such, the Nano Orbital Workshop has taken a new approach of only conducting modulated Exo-Brake research on missions deploying below ISS altitude to avoid collision avoidance restrictions, and is working to improve spacecraft identification and location reporting techniques to increase experiment durations. Upcoming TES mission will therefore use one of two classes of Exo-Brake depending on their target altitude, as to be described in the submission.

Marcus Murbach↗
NASA NTRS2002

Potential Operating Orbits for Fission Electric Propulsion Systems Driven by the SAFE-400

Safety must be ensured during all phases of space fission system design, development, fabrication, launch, operation, and shutdown. One potential space fission system application is fission electric propulsion (FEP), in which fission energy is converted into electricity and used to power high efficiency (Isp greater than 3000s) electric thrusters. For these types of systems it is important to determine which operational scenarios ensure safety while allowing maximum mission performance and flexibility. Space fission systems are essentially nonradioactive at launch, prior to extended operation at high power. Once high power operation begins, system radiological inventory steadily increases as fission products build up. For a given fission product isotope, the maximum radiological inventory is typically achieved once the system has operated for a length of time equivalent to several half-lives. After that time, the isotope decays at the same rate it is produced, and no further inventory builds in. For an FEP mission beginning in Earth orbit, altitude and orbital lifetime increase as the propulsion system operates. Two simultaneous effects of fission propulsion system operation are thus (1) increasing fission product inventory and (2) increasing orbital lifetime. Phrased differently, as fission products build up, more time is required for the fission products to naturally convert back into non-radioactive isotopes. Simultaneously, as fission products build up, orbital lifetime increases, providing more time for the fission products to naturally convert back into non-radioactive isotopes. Operational constraints required to ensure safety can thus be quantified.

Houts, Mike↗
NASA NTRS2002

Potential Operating Orbits for the SAFE-400

Safety must be ensured during all phases of space fission system design, development, fabrication, launch, operation, and shutdown. One potential space fission system application is fission electric propulsion (FEP), in which fission energy is converted into electricity and used to power high efficiency (Isp is greater than 3000s) electric thrusters. For these types of systems it is important to determine which operational scenarios ensure safety while allowing maximum mission performance and flexibility. Space fission systems are essentially non-radioactive at launch, prior to extended operation at high power. Once high power operation begins, system radiological inventory steadily increases as fission products build up. For a given fission product isotope, the maximum radiological inventory is typically achieved once the system has operated for a length of time equivalent to several half-lives. After that time, the isotope decays at the same rate it is produced, and no further inventory builds in. For an FEP mission beginning in Earth orbit, altitude and orbital lifetime increase as the propulsion system operates. Two simultaneous effects of fission propulsion system operation are thus (1) increasing fission product inventory and (2) increasing orbital lifetime. Phrased differently, as fission products build up, more time is required for the fission products to naturally convert back into non-radioactive isotopes. Simultaneously, as fission products build up, orbital lifetime increases, providing more time for the fission products to naturally convert back into non-radioactive isotopes. Operational constraints required to ensure safety can thus be quantified.

Houts, Mike↗
NASA NTRS2017

Drag De-Orbit Device: A New Standard Re-Entry Actuator for CubeSats

With the advent of CubeSats, research in Low Earth Orbit (LEO) becomes possible for universities and small research groups. Only a handful of launch sites can be used, due to geographical and political restrictions. As a result, common orbits in LEO are becoming crowded due to the additional launches made possible by low-cost access to space. CubeSat design principles require a maximum of a 25-year orbital lifetime in an effort to reduce the total number of spacecraft in orbit at any time. Additionally, since debris may survive re-entry, it is ideal to de-orbit spacecraft over unpopulated areas to prevent casualties. The Drag Deorbit Device (D3) is a self-contained targeted re-entry subsystem intended for CubeSats. By varying the cross-wind area, the atmospheric drag can be varied in such a way as to produce desired maneuvers. The D3 is intended to be used to remove spacecraft from orbit to reach a desired target interface point. Additionally, attitude stabilization is performed by the D3 prior to deployment and can replace a traditional ADACS on many missions.This paper presents the hardware used in the D3 and operation details. Four stepper-driven, repeatedly retractable booms are used to modify the cross-wind area of the D3 and attached spacecraft. Five magnetorquers (solenoids) over three axes are used to damp rotational velocity. This system is expected to be used to improve mission flexibility and allow additional launches by reducing the orbital lifetime of spacecraft.The D3 can be used to effect a re-entry to any target interface point, with the orbital inclination limiting the maximum latitude. In the chance that the main spacecraft fails, a timer will automatically deploy the booms fully, ensuring the spacecraft will at the minimum reenter the atmosphere in the minimum possible time, although not necessarily at the desired target interface point. Although this does not reduce the risk of casualties, the 25-year lifetime limit is still respected, allowing a reduction of the risk associated with a hardware failure.

reentry↗
NASA NTRS2017

Drag De-Orbit Device: A New Standard Re-Entry Actuator for CubeSats

With the advent of CubeSats, research in Low Earth Orbit (LEO) becomes possible for universities and small research groups. Only a handful of launch sites can be used, due to geographical and political restrictions. As a result, common orbits in LEO are becoming crowded due to the additional launches made possible by low-cost access to space. CubeSat design principles require a maximum of a 25-year orbital lifetime in an effort to reduce the total number of spacecraft in orbit at any time. Additionally, since debris may survive re-entry, it is ideal to de-orbit spacecraft over unpopulated areas to prevent casualties. The Drag Deorbit Device (D3) is a self-contained targeted re-entry subsystem intended for CubeSats. By varying the cross-wind area, the atmospheric drag can be varied in such a way as to produce desired maneuvers. The D3 is intended to be used to remove spacecraft from orbit to reach a desired target interface point. Additionally, attitude stabilization is performed by the D3 prior to deployment and can replace a traditional ADACS on many missions.This paper presents the hardware used in the D3 and operation details. Four stepper-driven, repeatedly retractable booms are used to modify the cross-wind area of the D3 and attached spacecraft. Five magnetorquers (solenoids) over three axes are used to damp rotational velocity. This system is expected to be used to improve mission flexibility and allow additional launches by reducing the orbital lifetime of spacecraft.The D3 can be used to effect a re-entry to any target interface point, with the orbital inclination limiting the maximum latitude. In the chance that the main spacecraft fails, a timer will automatically deploy the booms fully, ensuring the spacecraft will at the minimum reenter the atmosphere in the minimum possible time, although not necessarily at the desired target interface point. Although this does not reduce the risk of casualties, the 25-year lifetime limit is still respected, allowing a reduction of the risk associated with a hardware failure.

reentry↗
NASA NTRS1986

Nuclear reactor power for a space-based radar. SP-100 project

A space-based radar mission and spacecraft, using a 300 kWe nuclear reactor power system, has been examined, with emphasis on aspects affecting the power system. The radar antenna is a horizontal planar array, 32 X 64 m. The orbit is at 61 deg, 1088 km. The mass of the antenna with support structure is 42,000 kg; of the nuclear reactor power system, 8,300 kg; of the whole spacecraft about 51,000 kg, necessitating multiple launches and orbital assembly. The assembly orbit is at 57 deg, 400 km, high enough to provide the orbital lifetime needed for orbital assembly. The selected scenario uses six Shuttle launches to bring the spacecraft and a Centaur G upper-stage vehicle to assembly orbit. After assembly, the Centaur places the spacecraft in operational orbit, where it is deployed on radio command, the power system started, and the spacecraft becomes operational. Electric propulsion is an alternative and allows deployment in assembly orbit, but introduces a question of nuclear safety.

Bloomfield, Harvey↗
NASA NTRSDate not supplied

Eclipse Mitigation Strategies in P/2 Lunar Resonant Orbits

Lunar resonant orbits present several important challenges in terms of trajectory design, insertion, stability and eclipse mitigation. For some mission concepts, the placement of the initial orbital elements is fundamental to overcoming show-stopping performance values such as eclipse duration and orbit lifetime. Lunar resonant orbits need to maintain specific alignments in order to avoid undesired flybys or even recontact with Earth or GEO satellites. This paper presents methods to mitigate long eclipses while preserving achievable lunar resonant orbits that persist in the long term.

Eclipse Mitigation↗
NASA NTRS1992

Mass spectrometer use in a large chamber

The early satellites were somewhat insensitive to contamination produced during the construction and testing phases. The On-Orbit lifetime was such that contamination effects went either unnoticed or unrecognized. With today's On-Orbit lifetimes approaching 10+ years, contamination has become a paramount concern. The scientific payloads have increased in complexity and sensitivity. The ability to clean a contaminated sensor has greatly diminished. This requires better pumping systems and methods for improved monitoring. The conversion from diffusion pumped thermal vacuum chambers to cryo pumped chambers with the use of Misner traps and selective cold traps has reduced contamination. Witness samples supply a record of the condensates that remain after a testing cycle, but impart no knowledge of the contaminant migration during the cycle that may be a month in duration. Due to a customer's request that mass spectrometry be used during the testing of their spacecraft, a consultant was contracted to install a mass spectrometer to determine the feasibility of the instrument. The equipment and methodology described will start with the original system and its evolution to GE's present system.

Chuvala, Tom↗
NASA NTRS2006

The Disposal of Spacecraft and Launch Vehicle Stages in Low Earth Orbit

As a result of the increasing number of debris in low Earth orbit (LEO), numerous national and international orbital debris mitigation guidelines recommend the removal of spacecraft and launch vehicle stages from LEO within 25 years after mission termination. The primary purpose of this action is to enhance space safety by significantly limiting the potential of future accidental collisions resulting in the creation of large numbers of new orbital debris. Likewise, the passivation of these objects, i.e., the removal of residual stored energies, while they remain in orbit is important to prevent the generation of debris via self-induced explosions. Characteristics and trends in the growth of the derelict spacecraft and launch vehicle stage populations in LEO are examined. Depending upon the final operational altitude of the vehicle, achieving the goal of orbital lifetime reduction can influence the design and deployment philosophy of a new space system. Some spacecraft and launch vehicle stages have combined their end-of-mission passivation operations with maneuvers to vacate long-lived orbital regimes. Perhaps the most dramatic demonstration of this type occurred in 2006 when a U.S. Delta IV second stage executed an unprecedented controlled-reentry maneuver from a circular orbit at an altitude near 850 km. For space systems in orbital regimes near the upper regions of LEO (i.e., between 1400 km and 2000 km altitude), maneuvers to place the vehicle above LEO might be more attractive than attempting to ensure an atmospheric reentry within 25 years, and at least one space system operator has selected this option. In some cases, careful consideration of natural orbital perturbations can also lead to reduced orbital lifetimes, although new launch constraints might need to be imposed.

Johnson, Nicholas L.↗
NASA NTRS1974

Helium in the Martian atmosphere - Thermal loss considerations

Helium concentrations in the Martian atmosphere are estimated assuming that the helium production on Mars (comparable to its production on earth) via the radioactive decay of uranium and thorium is in steady state equilibrium with its thermal escape. Although nonthermal losses would tend to reduce the estimated concentrations, these concentrations are not necessarily an upper limit since higher production rates and/or a possibly lower effective exospheric temperature over the solar activity cycle could increase them to even higher values. The computed helium concentration at the Martian exobase (200 km) is 8 million atoms/cu cm. Through the lower exosphere, the computed helium concentrations are 30-200 times greater than the Mariner-measured atomic hydrogen concentrations. It follows that helium may be the predominant constituent in the Martian lower exosphere and may well control the orbital lifetime of Mars-orbiting spacecraft.

Levine, J. S.↗
NASA NTRS1984

Long-lifetime Martian orbit selection using a time-dependent model of the Martian atmosphere

A mathematical model of the time-dependent Martian atmosphere has been developed in order to accurately calculate the effects of aerodynamic drag on a low altitude satellite. The time-dependent properties of the model include solar activity effects, dust storm effects, seasonal and diurnal variations, and annual motion effects. Position effects are accounted for through Martian latitude and longitude. Expected values of mass density, temperature, scale height, and the estimated standard deviation of the mass density are provided. An example of the use of the model in selecting an orbit for the Mars Geochemical/Climatology Orbiter is given.

Culp, R. D.↗
NASA NTRS1986

Measured ground performance and predicted orbital performance of the superfluid helium dewar for the Cosmic Background Explorer

A critical component of the Cosmic Background Explorer observatory, which is to be lifted to orbit in 1988, is the 650-l superfluid He dewar housing a far-IR absolute spectrophotometer and a diffuse IR background experiment. Attention is presently given to the results of a four-month-long test program encompassing dewar filling verification, vibration characteristics, thermal performance over orbital lifetime, and aperture cover ejection behavior. No significant flaws have been noted; the orbital cryogen lifetime is projected to be 14 months.

Hopkins, Richard A.↗
NASA NTRS2008

Space Debris Symposium (A6.) Measurements and Space Surveillance (1.): Measurements of the Small Particle Debris Cloud from the 11 January, 2007 Chinese Anti-satellite Test

On January 11, 2007, the Chinese military conducted a test of an anti-satellite (ASAT) system, destroying their own Fengyun-1C spacecraft with an interceptor missile. The resulting hypervelocity collision created an unprecedented number of tracked debris - more than 2500 objects. These objects represent only those large enough for the US Space Surveillance Network (SSN) to track - typically objects larger than about 5-10 cm in diameter. There are expected to be even more debris objects at sizes too small to be seen and tracked by the SSN. Because of the altitude of the target satellite (865 x 845 km orbit), many of the debris are expected to have long orbital lifetimes and contribute to the orbital debris environment for decades to come. In the days and weeks following the ASAT test, NASA was able to use Lincoln Laboratory s Haystack radar on several occasions to observe portions of the ASAT debris cloud. Haystack has the capability of detecting objects down to less than one centimeter in diameter, and a large number of centimeter-sized particles corresponding to the ASAT cloud were clearly seen in the data. While Haystack cannot track these objects, the statistical sampling procedures NASA uses can give an accurate statistical picture of the characteristics of the debris from a breakup event. For years computer models based on data from ground hypervelocity collision tests (e.g., the SOCIT test) and orbital collision experiments (e.g., the P-78 and Delta-180 on-orbit collisions) have been used to predict the extent and characteristics of such hypervelocity collision debris clouds, but until now there have not been good ways to verify these models in the centimeter size regime. It is believed that unplanned collisions of objects in space similar to ASAT tests will drive the long-term future evolution of the debris environment in near-Earth space. Therefore, the Chinese ASAT test provides an excellent opportunity to test the models used to predict the future debris environment. For this study, Haystack detection events are compared to model predictions to test the model assumptions, including debris size distribution, velocity distribution, and assumptions about momentum transfer between the target and interceptor. In this paper we will present the results of these and other measurements on the size and extent of collisional breakup debris clouds.

Matney, Mark J.↗