National Aeronautics And Space Administration Orbital Debris

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National Aeronautics and Space AdministrationOrbitalDebrisQuarterly NewsVolume 24, Issue 3August 2020International Space Station Maneuversto Avoid DebrisInside.Second Fragmentationof Fregat Upper StageDebris2DAS v3.1 Release3Developmentof ExperimentalHypervelocity ImpactCapabilities with NonSpherical Projectiles 4Short-TermSatellite BreakupRisk AssessmentModel Process7Two Recent BreakupEvents Updated11Space Missions andSatellite Box Score 14A publication of theNASA Orbital DebrisProgram Office (ODPO)The International Space Station (ISS) conductedtwo maneuvers to avoid potential collisions withlarge debris tracked by the U.S. Space Command(USSPACECOM) Space Surveillance Network (SSN)on 19 April and 3 July.The maneuver in April was triggered by ahigh-risk conjunction with a breakup fragmentgenerated from the Fengyun-1C (FY-1C) anti-satellitetest conducted by China in 2007. That fragment hasan International Designator of 1999-025BNN and anSSN Catalog ID of 31280. Because of the conjunction’stiming, the ISS Program used the opportunity of aplanned deboost maneuver (to set up the proper orbitconfiguration for upcoming ISS visiting vehicles) toalso avoid the high-risk conjunction with the FY-1Cdebris.The July maneuver was to avoid an object withInternational Designator 1987-079AG and SSN#27923, which was generated from the explosion of aSOZ (Sistema Obespecheniya Zapuska, “Launch SupportSystem”) ullage motor, or SL-12 auxiliary motor, in2003. A total of 42 fragments from the explosion ofthat SOZ motor (International Designator 1987-079H,SSN# 18375) are large enough to catalog, includingobject 27923. Due to a design flaw, more than 50 SOZullage-motor explosions have been documented so far.The ISS has conducted a total of 27 collisionavoidance maneuvers since 1999. The avoided objectsincluded two FY-1C debris and five debris from the2009 collision between Cosmos 2251 and Iridium 33(two from the former and three from the latter). Inaddition, there was a high-risk conjunction betweenthe ISS and NASA’s operational Global PrecipitationMeasurement (GPM) spacecraft in 2017, and the GPMspacecraft maneuvered to mitigate the risk of collision.The history of the ISS collision avoidancemaneuvers, the number of SSN-tracked, ISSorbit-crossing objects, and the solar activity forF10.7 cm flux are shown in the figure on page 2. Thenumber of the SSN-tracked, ISS-orbit-crossing objectsis influenced by several factors, including space trafficand debris population, solar activity, and the altitudeof the ISS, which varied between approximately320 km and 430 km over the years. The frequency ofthe ISS collision avoidance maneuvers is affected by theISS conjunction risk mitigation protocol, the numberof the ISS-orbit-crossing objects, and the accuracyof conjunction assessments. The latter depends onsensor capability, data processing and analysis, andsolar activity. Under the leadership of the 18th SpaceControl Squadron, the accuracy of object tracking andconjunction assessments has improved in recent years.Orbital debris has been identified as a majorsafety risk for the ISS with respect to loss of missionand loss of crew. The U.S. modules on the ISS arewell-protected against impacts by small orbital debris.The protective shields are effective against orbitaldebris about 1 cm and smaller. The objects shown ascircles in the figure are large enough to be tracked bythe SSN, which are objects approximately 10 cm andlarger. Collision risk against such large objects can bemitigated by conjunction assessments and collisionavoidance maneuvers when necessary. Therefore, thebiggest threat to the U.S. modules on the ISS comesfrom orbital debris between 1 cm and 10 cm. Sincethe orbital debris population follows a power-law sizedistribution, meaning there is a higher number of smalldebris than large debris, the risk to the U.S. modulesis actually driven by orbital debris in the 1 cm–2 cmrange. The number of such small debris crossing theISS orbit is about 20 times greater than the blue circlesshown in the figure. Other non-U.S. modules on theISS are not as well protected against small orbitaldebris impacts. The risk to those modules is driven bysub-centimeter orbital debris. see figure on page 2

Orbital Debris Quarterly NewsISS Maneuverscontinued from page 2ISS‐Orbit‐Crossing Tracked Objects, ISS Collision Avoidance Maneuvers, & Solar Flux12ISS‐orbit‐crossing tracked objectsISS collision avoidance maneuvers100010Solar F10.7 Daily Flux800860064004200201995Number of ISS Collision Avoidance Maneuvers[Number Of Objects] or [F10.7 Daily Flux (sfu)]12000200020052010201520202025TimeNumber of the ISS-orbit-crossing objects tracked by the SSN (blue circles), the solar F10.7 daily flux (black dots), and the history of the ISS collision avoidance maneuvers(orange histogram).Second Fragmentation of Fregat Upper Stage DebrisA debris object associated with the launch of Russia’s Spektr-R radioastronomy satellite fragmented on 8 May 2020 between 0402 GMT and0551 GMT. The object (International Designator 2011-037B, U.S. SpaceCommand Space Surveillance Network [SSN] catalog number 37756) hadexperienced a prior breakup event, of unknown cause, in August 2015.Previously described in the SSN catalog as “SL-23 DEB,” the objectis now identified as “FREGAT DEB (TANK)”. The Lavochkin Fregat-SBused in this launch is based on the Fregat upper stage but adds a toroidalhypergolic fuel/oxidizer tank, the sbrasyvaemye blok bakov (SBB) andvariously referred to as the jettisoned tanks unit (JTU) or block (JTB).The reader is referred to the previous ODQN news article (ODQN,vol. 20, Issues 1 & 2 joint issue, pp. 2-3*) for detailed illustrations anda description of the physical attributes of the vehicle, including storedenergy and possible failure modes.At the time of the May 2020 event, the SBB was in a 3606 x 422 kmaltitude, 51.5 inclination orbit. It had been on-orbit for approximately4.0 years and 8.8 years at the respective times of the two events.2A total of 24 fragments were initially tracked from the August2015 event. As of 5 June 2020, however, none had been confirmed andcataloged. The May 2020 event produced approximately 65 trackeddebris of which 36, including the parent body, had entered the catalog by5 June. The 2011-037 debris ensemble of the parent body (piece tag B)and new debris (piece tags H through AT, inclusive) are portrayed in thefigure’s Gabbard diagram (shown on pg. 3).While the presence of residual stored energy, particularly thehypergolic propellants, offers a likely root cause of this energetic event,uncertainties associated with actual construction details render the actualcause unknown at this time. * Note: The ODQN article’s second reference, “Propulsion System for Delivering“Phobos-Grunt” Spacecraft on Phobos Surface,” is now identified as havingbeen authored by Yu.G. Stekolshchikov, S.S. Stepanov, L.G. Alexandrov,and V.P. Makarov of the Lavochkin Association.see figure on page 3

https://orbitaldebris.jsc.nasa.govFregat Fragmentationcontinued from page 22011‐037B Debris Cloud, 5 June 2020 epoch60005000Altitude [km]4000Debris apogee altitude [km]3000Debris perigee altitude [km]Parent apogee altitude [km]2000Parent perigee altitude [km]10000115120125130135140145150Period [min]A Gabbard diagram of the 8 May 2020 2011-037B event. Approximate epoch is 5 June 2020.The maximum change in period is approximately 20 minutes, and the maximumchange in inclination is 0.21 .Debris Assessment Software Version 3.1 ReleaseThe Orbital Debris Program Office has released version 3.1 of theDebris Assessment Software (DAS), replacing the prior October 2019release of DAS 3.0. The updated version provides data that can verifycompliance of a spacecraft, upper stage, and/or payload with NASA’srequirements for limiting debris generation, spacecraft vulnerability,postmission lifetime, and entry safety.This release incorporates an update to the Orbital DebrisEngineering Model (ORDEM), version 3.1, which was unveiled inODQN, vol. 24, issue 1, p. 3. Another change to DAS is the removalof the Grün meteoroid flux model. To calculate penetration risk frommeteoroids, users should consult the NASA Meteoroid EnvironmentOffice and Hypervelocity Impact Technology teams for assessments.Successful verification of a design in DAS demonstrates compliancewith NASA debris mitigation requirements. Historically, DAS analysishas proven acceptable in meeting compliance requirements of manyother agencies in the U.S. and around the world. It does not address theinherent design reliability facets of NASA requirements, but addressesall Earth-related orbital debris requirements that make up the bulk ofthe requirements in the NASA Technical Standard 8719.14B.For new users, DAS is available for download, by permission only,and requires that an application be completed via the NASA SoftwareCatalog. To begin the process, click on the Request Now! button in thecatalog at https://software.nasa.gov/software/MSC-26690-1 .Users who have already completed the software request processfor earlier versions of DAS 3.x do not need to reapply for DAS 3.1.Simply go to your existing account on the NASA Software portal anddownload the latest installer. Approval for DAS is on a per project basis:approval encompasses activities and personnel working within theproject scope identified in the application. 3

Orbital Debris Quarterly NewsPROJECT REVIEWThe Development of Experimental HypervelocityImpact Capabilities with Non-Spherical ProjectilesJ. MILLER, B. DAVIS, T. JUDD, AND R. MCCANDLESS(HYPERVELOCITY IMPACT TECHNOLOGY GROUP)A. DELGADO, D. HENDERSON, A. PARDO,D. RODRIGUEZ, AND M. SANDY(REMOTE HYPERVELOCITY TEST LABORATORY)In 2014 a laboratory impact experiment, DebriSat, was conductedat Arnold Engineering Development Center. A still frame image seriesof the experiment is shown in Figure 1 [1, 2]. DebriSat was designedto replicate a catastrophic impact to update the NASA satellite breakupmodel and to support the development of a fragment shape distributionfor the future Orbital Debris Engineering Model (ORDEM 4.0), whichwill improve the fidelity of orbital debris impact risk assessments[2, 3]. To this end, DebriSat was built with many modern materialsincluding structural panels of carbon-fiber reinforced polymer (CFRP).Subsequent to the experiment, DebriSat fragments were extractedfrom porous catcher foam panels that minimized secondary damage toFigure 1. The DebriSat experiment subjected a representative modern satellite to ahypervelocity impact (top) of a 570 g-aluminum projectile at 6.8 km/s, which resultedin the catastrophic break-up of the satellite. A large database of debris was capturedand is being categorized with a large portion of the debris being (bottom) carbon-fiberreinforced polymer with shapes from large and flat “flake-like” objects to long and thin“needle-like” objects. Credit: (top) Arnold Engineering Development Complex/UnitedStates Air Force; (bottom) University of Florida, Dept. of Mechanical & AerospaceEngineering, Space Systems Group.Beta Cloth/Mylar (0.026 g/cm²)3.2 mmFiberglass (0.030 g/cm²)0.2 mm Al (0.054 g/cm²)3.2 mmFiberglass (0.030 g/cm²)MLI (0.036 g/cm²)15 mm0.5 mm Al 6061-T6 (0.135 g/cm²)2.0 mm Al 5456-O (0.53 g/cm²)50 mm1.0 mm Al 2024-T4 (0.28 g/cm²)the debris during the capture process [3, 4]. Thus far, one of the keyobservations is that CFRP fragments represent a large fraction of thecollected debris and that these fragments tend to be thin, “flake-like”structures or long, “needle-like” structures, as illustrated in Figure 1;whereas, debris with nearly equal dimensions is less prevalent. Ascurrent ballistic-limit models for shields are based upon sphericalimpacting particles [5], the DebriSat experiment has pointed to a missingcomponent in the current approach to ballistic modeling that must beconsidered. To improve risk assessments of spacecraft design reliabilityand survivability, refined, broad-ranging, non-spherical ballistic limitequations are needed to prepare for the future Bumper implementationof ORDEM 4.0. This effort is funded by the NASA HQ Office of Safetyand Mission Assurance (OSMA).While numerous shield types are currently in use for impactmitigation from orbital debris and meteoroids, the most common shieldin use is the double-wall shield commonly known as a Whipple shield[6]. This shield achieves a high level of ballistic performance for minimalweight because the stresses induced in a projectile during impact are farabove the stresses the solid particle can withstand, resulting in a break-upof the particle. In the Whipple shield approach, an empty volumebetween the two walls of the shield provides a space for the debris cloudto expand, resulting in a distributed impact on the second shield-wall;however, even with the increased performance of this design, the shieldwall reaches a limit, which is the ballistic-limit of the shield-system [7].The pure, all-metal, Whipple shield is less prevalent in deploymentdue to operational thermal environments [5]. A previous ODQN article,(ODQN, vol. 22, issue 4, November 2018, pp. 2-4), [8] developed anumerical simulation model of a representative International SpaceStation (ISS) shield with the material configuration shown in Figure 2.This shield carries a higher risk of penetration than most other ISS shieldsdue to its lower ballistic capability, and therefore, its performance is ofparticular interest to the operation of the ISS.This numerical simulation model was executed over a broad range ofimpact conditions to capture the effect of the non-spherical nature of thedebris collected from DebriSat. However, accounting for geometricproperties, like material thicknesses, separation, and orientation, is notsufficient to address potential concerns with material constitutivepropertiesinanumerical simulationmodel. To ensure thatmodeling of the materialconstitutive propertieshas been performedproperly, a collectionof impact validationdata is needed forcomparison to full-scalenumerical simulations ofthe impact event.To this end, theHypervelocity ImpactTechnology group at theNASA Johnson SpaceFigure 2. Representative Whipple shield with external thermal blanket (left) shield configuration schematic (layers scaled by mass; separations toscale), and (right) profile image of representative experimental article.4continued on page 5

https://orbitaldebris.jsc.nasa.govImpact Capabilitiescontinued from page 4Center in Houston, Texas, and the Remote Hypervelocity Test LaboratoryThe cameras, located 45 down from the top of the target chamber,at the NASA White Sands Test Facility in Las Cruces, New Mexico, teamed are focused on an alignment fixture (Figure 4, top, right-hand image)up to obtain 11 representative impacts of cylindrical CFRP projectilescontinued on page 6at about seven kilometers per second into theshield shown in Figure 2 [9]. These experimentsperformed two functions: first, they allowed thedevelopment of manufacturing techniques, launchconditions, and diagnostic techniques necessaryto achieve ballistic data for cylindrical projectilesat orbital speeds, and second, they provideddata for comparison to numerical simulationsfor the popular Whipple shield system. Thecylindrical projectiles fell into three generalcategories: “flake-like,” with ratio of length (L) todiameter (D), L:D, of approximately 1:3, where arepresentative example is shown in Figure 3 (lefthand side), spherical mass-equivalent with L:D of2:3 shown in Figure 3 (center), and “needle-like”with an L:D ratio of 3:1 shown in Figure 3 (righthand side).The projectiles have been fabricated fromboth axially extruded Torayca T700 fiber systemsin a rod-stock form with a high-performance Figure 3. Side and top images of representative cylindrical, CFRP projectiles for (left-hand) “flake-like” at L:D 1:3(actual L:D range of 1:2.5-1:5, determined by fixed thickness of stock material and variable diameter; L:D 1:4.8Bisphenol A epoxy resin binder [10], and a in this case); (center) spherical mass-equivalent at L:D 2:3; and (right-hand) “needle-like” at L:D 3:1.plane-woven, 1/16 inch-thick, sheet-stock ofUltra-Strength, Lightweight Carbon Fiber [11].The projectile geometry for the “needle-like”and spherical mass-equivalent projectiles aredetermined by the initial diameter of the parentrod-stock material and are cut to length with arotating very-fine-grit, diamond-cutoff wheelmounted on the tail stock of a counter-rotatinglathe. The projectile geometry for the “flakelike” projectiles is derived from the sheet-stockmaterial and is cut with a high-pressure, waterjet cutter to the required diameter. All projectiledimensions are characterized using a KeyenceVHX-5000 Digital Microscope with VH-Z20Robjective to a 25 µm resolution. All threeprojectile geometries are accelerated within aseparable sabot by a 0.50 caliber, two-stage, lightgas gun with the cylindrical axis of the projectilein the direction of the barrel.While all projectiles are accelerated withtheir cylindrical axis pointed toward the target,the release from the carrying sabot and the flightwithin the target chamber results in the potentialfor an arbitrary rotation of the projectile. Toaddress this potential rotation, two framesynchronized Shimadzu HyperVision HPV-X Figure 4. (top images) Pre-impact imagery of the projectile relies on a pair of frame-synchronized, ultrahigh-speedcameras with Nikon AF Zoom-NIKKOR 80- Shimadzu HyperVision HPV X cameras located outside the target chamber (left-hand image) to image the target200 mm Extra-low Dispersion lenses image the (center image, the dual attachment fixture in detail). The cameras are aligned orthogonal to each other and focusedprojectile at 2 million frames per second as the on the flight path of the projectile prior to each shot by an in situ alignment tool consisting of five, precisely located,projectile approaches the target. For maximum all-thread rods (right-hand image). The two, 2 million-frame-per-second cameras are capable of capturing aboutlight collection, the objective lenses have been set 40 frames of the projectiles moving at 7 km/s as they move toward the target over their field of view.at f/2.8.These two cameras are placed orthogonal (bottom image) These images of the projectile just prior to the moment of impact enable the calculation of theto each other outside of the target chamber as orientation of the projectile with respect to its velocity vector. In these backlit images, the shadow of the projectileis about 10 µs from the target on the far right and the visible radiation from the bow shock wave generated by theseen in Figure 4 (left-hand image).projectile is also visible in the image.5

Orbital Debris Quarterly NewsImpact Capabilitiescontinued from page 5that is placed immediately in front of the target. The alignment fixtureprovides a reference to adjust the cameras so that they are orthogonal,has features for focus adjustment at the expected flight path, and providesa spatial absolute reference for scaling the camera images at the expectedposition of the projectile in flight.A secondary target frame, also shown in Figure 4 (top, center image),directly interfaces with the target chamber and holds the primary targetof Figure 2 and alignment tool for each shot. This dual frame approachallows the rapid and repeatable change-out of targets between successiveshots without altering the camera alignment. The secondary target framealso provides background screens to each of the Shimadzu cameras.These background screens reflect the light from the illuminationsource at the top of the target chamber directly back to the cameras,allowing an approximately flat backlighting of the projectile as itapproaches the primary target, as shown in the scaled images of Figure 4(bottom image). From these two orthogonal views, angles between thecylinder’s axis and the velocity vector can be measured for each view.These orthogonal rotation angles can then be used to calc

While numerous shield types are currently in use for impact mitigation from orbital debris and meteoroids, the most common shield in use is the double-wall shield commonly known as a Whipple shield [6]. This shield achieves a high level of ballistic performance for minimal weight because the stresses induced in a projectile during impact are far

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