Results for “TUMBLING” — Page 10

NASA NTRS2004

Genesis failure investigation report : JPL Failure Review Board, Avionics Sub-Team

On January 7, 2001, the Genesis spacecraft lifted off from Cape Canaveral. Its mission was to collect solar wind samples and return those samples to Earth for detailed analysis by scientists. The mission proceeded successfully for three-and-a-half years. On September 8, 2004, the spacecraft approached Earth, pointed the Sample Return Capsule (SRC) at its entry target, and then fired pyros that jettisoned the SRC. The SRC carried the valuable samples collected over the prior 29 months. The SRC also contained the requisite hardware (mechanisms, parachutes, and electronics) to manage the process of entry, descent, and landing (EDL). After entering Earthas atmosphere, the SRC was expected to open a drogue parachute. This should have been followed by a pyro event to release the drogue chute, and then by a pyro event to deploy the main parachute at an approximate elevation of 6.7 kilometers. As the SRC descended to the Utah landing site, helicopters were in position to capture the SRC before the capsule touched down. On September 8, 2004, observers of the SRCas triumphant return became concerned as the NASA announcer fell silent, and then became even more alarmed as they watched the spacecraft tumble as it streaked across the sky. Long-distance cameras clearly showed that the drogue parachute had not deployed properly. On September 9, 2004, General Eugene Tattini, Deputy Director of the Jet Propulsion Laboratory formed a Failure Review Board (FRB). This board was charged with investigating the cause of the Genesis mishap in close concert with the NASA Mishap Investigation Board (MIB). The JPL-FRB was populated with experts from within and external to the Jet Propulsion Laboratory. The JPL-FRB participated with the NASA-MIB through all phases of the investigation, working jointly and concurrently as one team to discover the facts of the mishap.

solar wind↗

NASA NTRS1964

Dynamic Model Investigation of the Landing Characteristics of a Manned Spacecraft

Investigations were made to study the water-landing and certain grounds-surface landing characteristics of a Gemini spacecraft model. The water landing experiments were made by simulating paraglider and parachute letdowns with two 1/6- scale model configurations. Parameters included various combinations of attitude, horizontal speed, vertical speed, and landing skids extended and retracted. Investigations were made in calm water and in waves. The paraglider landings at horizontal speeds of 63 feet per second (19.8 m/sec) which resulted in a noseover or tumbling shortly after initial water contact. The maximum longitudinal acceleration of the model in calm water was about 14g units, and the maximum angular acceleration was 66 radians per second squared. In the parachute landings with the heat shield forward, the model skidded along the water surface on the heat shield. Parachute landings with the small end forward resulted in behavior similar to that of the paraglider landings. The ground-surface landings were made with a 1/3-scale model by simulating a parachute letdown with braking rockets, which were fired prior to touchdown to dissipate vertical velocity. In these landings, control of timing and aligning the rockets on the model was very critical, and violent behavior resulted when either rocket alignment or timing was in error. In the landings that were correctly controlled, the model either remained upright or slowly rolled over on its side.

EXPERIMENTS↗

NASA NTRS1963

Dynamic Model Investigation of a 1/20 Scale Gemini Spacecraft in the Langley Spin Tunnel

The investigation was conducted in the Langley spin tunnel. The tunnel is an atmospheric wind tunnel with a vertically rising airstream in the test section and a maximum airspeed of approximately 90 feet per second. For this investigation, the model was hand launched into the vertically rising airstream. At times the model, both with and without a drogue parachute, was launched gently with as little disturbance as possible to determine what motions of the spacecraft were self-excited. At other times, the spacecraft with pre-deployed drogue parachute was launched into various spinning motions to determine the effectiveness of the drogue parachute in terminating these spinning motions. During drogue-parachute deployment tests, the spacecraft was launched into various spinning and tumbling motions and the drogue parachute was deployed. The motions of the model were photographed with a motion-picture camera, and some of the film records were read to obtain typical time histories of the model motion. The angles of attack indicated in the time histories presented are believed to be accurate within +/-1 degree. The mass and dimensional characteristics of the dynamic model are believed to be measured to an accuracy of: +/-1 percent for the weight, +/-1 percent for z(sub cg)/d, +/-15 percent for x (sub cg), and +/-5 percent for the moments of inertia. The towline and bridle-line lengths were simulated to an accuracy of +/-1 foot full scale.

Lee, Henry A.↗

NASA NTRS2007

Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Cepollina, Frank J.↗

NASA NTRS2007

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Cepollina, Frank J.↗

NASA NTRS2005

Reconstruction of the Genesis Entry

This paper provides an overview of the findings from a reconstruction analysis of the Genesis capsule entry. First, a comparison of the atmospheric properties (density and winds) encountered during the entry to the pre-entry profile is presented. The analysis that was performed on the video footage (obtained from the tracking stations at UTTR) during the descent is then described from which the Mach number at the onset of the capsule tumble was estimated following the failure of the drogue parachute deployment. Next, an assessment of the Genesis capsule aerodynamics that was extracted from the video footage is discussed, followed by a description of the capsule hypersonic attitude that must have occurred during the entry based on examination of the recovered capsule heatshield. Lastly, the entry trajectory reconstruction that was performed is presented.

Desai, Prasun N.↗

NASA NTRS2007

Automated Target Planning for FUSE Using the SOVA Algorithm

The SOVA algorithm was originally developed under the Resilient Systems and Operations Project of the Engineering for Complex Systems Program from NASA s Aerospace Technology Enterprise as a conceptual framework to support real-time autonomous system mission and contingency management. The algorithm and its software implementation were formulated for generic application to autonomous flight vehicle systems, and its efficacy was demonstrated by simulation within the problem domain of Unmanned Aerial Vehicle autonomous flight management. The approach itself is based upon the precept that autonomous decision making for a very complex system can be made tractable by distillation of the system state to a manageable set of strategic objectives (e.g. maintain power margin, maintain mission timeline, and et cetera), which if attended to, will result in a favorable outcome. From any given starting point, the attainability of the end-states resulting from a set of candidate decisions is assessed by propagating a system model forward in time while qualitatively mapping simulated states into margins on strategic objectives using fuzzy inference systems. The expected return value of each candidate decision is evaluated as the product of the assigned value of the end-state with the assessed attainability of the end-state. The candidate decision yielding the highest expected return value is selected for implementation; thus, the approach provides a software framework for intelligent autonomous risk management. The name adopted for the technique incorporates its essential elements: Strategic Objective Valuation and Attainability (SOVA). Maximum value of the approach is realized for systems where human intervention is unavailable in the timeframe within which critical control decisions must be made. The Far Ultraviolet Spectroscopic Explorer (FUSE) satellite, launched in 1999, has been collecting science data for eight years.[1] At its beginning of life, FUSE had six gyros in two IRUs and four reaction wheels. Over time through various failures, the satellite has been left with one reaction wheel on the vehicle skew axis and two gyros. To remain operational, a control scheme has been implemented using the magnetic torque rods and the remaining momentum wheel.[2] As a consequence, there are attitude regions where there is insufficient torque authority to overcome environmental disturbances (e.g. gravity gradient torques). The situation is further complicated by the fact that these attitude regions shift inertially with time as the spacecraft moves through earth s magnetic field during the course of its orbit. Under these conditions, the burden of planning targets and target-to-target slew maneuvers has increased significantly since the beginning of the mission.[3] Individual targets must be selected so that the magnetic field remains roughly aligned with the skew wheel axis to provide enough control authority to the other two orthogonal axes. If the field moves too far away from the skew axis, the lack of control authority allows environmental torques to pull the satellite away from the target and can potentially cause it to tumble. Slew maneuver planning must factor the stability of targets at the beginning and end, and the torque authority at all points along the slew. Due to the time varying magnetic field geometry relative to any two inertial targets, small modifications in slew maneuver timing can make large differences in the achievability of a maneuver.

Heatwole, Scott↗

NASA NTRS2007

Reconstruction of the Genesis Entry

An overview of the reconstruction analyses performed for the Genesis capsule entry is described. The results indicate that the actual entry prior to the drogue deployment failure was very close to the pre-entry predictions. The capsule landed 8.3 km south of the desired target at Utah Test and Training Range. Analysis on infrared video footage (obtained from the tracking stations) during the descent estimated the onset of the capsule tumble at Mach 0.9. Frequency analysis on the infrared video data indicates that the aerodynamics generated for the Genesis capsule reasonably predicted the drag and static stability. Observations of the heatshield support the pre-entry simulation estimates of a small hypersonic angles-of-attack, since there is very little, if any, charring of the shoulder region or the aftbody. Through this investigation, an overall assertion can be made that all the data gathered from the Genesis entry is consistent with flight performance that was close to the nominal preentry prediction. Consequently, the design principles and methodologies utilized for the flight dynamics, aerodynamics, and aerothermodynamics analyses have been corroborated.

Desai, Prasun N.↗

NASA NTRS2008

NanoSail-D: The First Flight Demonstration of Solar Sails for Nanosatellites

The NanoSail-D mission is currently scheduled for launch onboard a Falcon Launch Vehicle in the late June 2008 timeframe. The NanoSail-D, a CubeSat-class satellite, will consist of a sail subsystem stowed in a Cubesat 2U volume integrated with a CubeSat 1U volume bus provided by the NASA Ames Research Center (ARC). Shortly after deployment of the NanoSail-D from a Poly Picosatellite Orbital Deployer (P-POD) ejection system, the solar sail will deploy and mission operations will commence. This demonstration flight has two primary mission objectives: 1) to successfully stow and deploy the sail and 2) to demonstrate de-orbit functionality. Given a nearterm opportunity for launch, the project was met with the challenge of delivering the flight hardware in approximately six months, which required a significant constraint on flight system functionality. As a consequence, passive attitude stabilization will be achieved using permanent magnets to de-tumble and orient the body with the magnetic field lines and then rely on atmospheric drag to passively stabilize the sailcraft in an essentially maximum drag attitude. This paper will present an introduction to solar sail propulsion systems, overview the NanoSail-D spacecraft, describe the performance analysis for the passive attitude stabilization, and present a prediction of flight data results from the mission.

Whorton, Mark↗

NASA NTRS2008

Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Cepollina, Frank J.↗

NASA NTRS1969

America in Space: The First Decade - NASA Spacecraft

It is ten years since the National Aeronautics and Space Administration was created to explore space and to continue the American efforts that had already begun with the launch of Explorer 1 on January 31, 1958. Many changes have occurred since that tumbling, 31 -pound cylinder went into an Earth orbit. "NASA Spacecraft" represents one of the broad avenues selected by NASA as an approach to its objective of making widely known the progress that has taken place in its program of space exploration. This report is a vivid illustration of the changes that have occurred and the complexities that have developed. Here one finds descriptions of the present family of spacecraft some small, some large; some spinoriented, some accurately attitude-controlled; some manned, some automated; some in low orbits, some in trajectories to the Moon and the planets; some free in space until they expire, others commanded to return to the Earth or to land on the Moon

Source record↗

NASA NTRS2009

Real-Time Simulation of Ares I Launch Vehicle

The Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) has been developed for use by the Ares I launch vehicle System Integration Laboratory (SIL) at the Marshall Space Flight Center (MSFC). The primary purpose of the Ares SIL is to test the vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment to certify that the integrated system is prepared for flight. ARTEMIS has been designed to be the real-time software backbone to stimulate all required Ares components through high-fidelity simulation. ARTEMIS has been designed to take full advantage of the advances in underlying computational power now available to support HWIL testing. A modular real-time design relying on a fully distributed computing architecture has been achieved. Two fundamental requirements drove ARTEMIS to pursue the use of high-fidelity simulation models in a real-time environment. First, ARTEMIS must be used to test a man-rated integrated avionics hardware and software system, thus requiring a wide variety of nominal and off-nominal simulation capabilities to certify system robustness. The second driving requirement - derived from a nationwide review of current state-of-the-art HWIL facilities - was that preserving digital model fidelity significantly reduced overall vehicle lifecycle cost by reducing testing time for certification runs and increasing flight tempo through an expanded operational envelope. These two driving requirements necessitated the use of high-fidelity models throughout the ARTEMIS simulation. The nature of the Ares mission profile imposed a variety of additional requirements on the ARTEMIS simulation. The Ares I vehicle is composed of multiple elements, including the First Stage Solid Rocket Booster (SRB), the Upper Stage powered by the J- 2X engine, the Orion Crew Exploration Vehicle (CEV) which houses the crew, the Launch Abort System (LAS), and various secondary elements that separate from the vehicle. At launch, the integrated vehicle stack is composed of these stages, and throughout the mission, various elements separate from the integrated stack and tumble back towards the earth. ARTEMIS must be capable of simulating the integrated stack through the flight as well as propagating each individual element after separation. In addition, abort sequences can lead to other unique configurations of the integrated stack as the timing and sequence of the stage separations are altered.

Tobbe, Patrick↗

NASA NTRS2009

Photometric Studies of Orbital Debris at GEO

Orbital debris represents a significant and increasing risk to operational spacecraft. Here we report on photometric observations made in standard BVRI filters at the Cerro Tololo Inter-American Observatory (CTIO) in an effort to determine the physical characteristics of optically faint debris at geosynchronous Earth orbit (GEO). Our sample is taken from GEO objects discovered in a survey with the University of Michigan s 0.6-m Curtis-Schmidt telescope (known as MODEST, for Michigan Orbital DEbris Survey Telescope), and then followed up in real-time with the CTIO/SMARTS 0.9-m for orbits and photometry. For a sample of 50 objects, calibrated sequences in RB- V-I-R filters have been obtained with the CTIO/SMARTS 0.9-m. For objects that do not show large brightness variations, the colors are largely redder than solar in both B-R and R-I. The width of the color distribution may be intrinsic to the nature of the surfaces, but also could imply that we are seeing irregularly shaped objects and measuring the colors at different times with just one telescope. For irregularly shaped objects tumbling at unknown orientations and rates, such sequential filter measurements using one telescope are subject to large errors for interpretation. If all observations in all filters in a particular sequence are of the same surface at the same solar and viewing angles, then the colors are meaningful. Where this is not the case, interpretation of the observed colors is impossible. For a smaller sample of objects we have observed with synchronized CCD cameras on the two telescopes. The CTIO/SMARTS 0.9-m observes in B, and the Schmidt in R. The CCD cameras are electronically linked together so that the start time and duration of observations are both the same to better than 50 milliseconds. Now the observed B-R color is a true measure of the scattered illuminated area of the debris piece for that observation.

Seitzer, Patrick↗

NASA NTRS2009

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR, and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time as well as the ability to intervene using manual override to teleoperate the robot.

Cepollina, Frank J.↗

NASA NTRS2009

Method and associated apparatus for capturing, servicing, and de-orbiting earth satellites using robotics

This invention is a method and supporting apparatus for autonomously capturing, servicing and de-orbiting a free-flying spacecraft, such as a satellite, using robotics. The capture of the spacecraft includes the steps of optically seeking and ranging the satellite using LIDAR; and matching tumble rates, rendezvousing and berthing with the satellite. Servicing of the spacecraft may be done using supervised autonomy, which is allowing a robot to execute a sequence of instructions without intervention from a remote human-occupied location. These instructions may be packaged at the remote station in a script and uplinked to the robot for execution upon remote command giving authority to proceed. Alternately, the instructions may be generated by Artificial Intelligence (AI) logic onboard the robot. In either case, the remote operator maintains the ability to abort an instruction or script at any time, as well as the ability to intervene using manual override to teleoperate the robot.In one embodiment, a vehicle used for carrying out the method of this invention comprises an ejection module, which includes the robot, and a de-orbit module. Once servicing is completed by the robot, the ejection module separates from the de-orbit module, leaving the de-orbit module attached to the satellite for de-orbiting the same at a future time. Upon separation, the ejection module can either de-orbit itself or rendezvous with another satellite for servicing. The ability to de-orbit a spacecraft further allows the opportunity to direct the landing of the spent satellite in a safe location away from population centers, such as the ocean.

Cepollina, Frank J.↗

NASA NTRS2010

The Fifteen-Year Attitude History of the Wide Field Planetary Camera 2 Radiator and Collection Efficiencies for Micrometeoroids and Orbital Debris

An examination of the Hubble Space Telescope (HST) Wide Field Planetary Camera 2 (WFPC-2) radiator assembly was conducted at NASA Goddard Space Flight Center (GSFC) during the summer of 2009. Immediately apparent was a distinct biasing of the largest 45 impact features towards one side of the radiator, in contrast to an approximately uniform distribution of smaller impacts. Such a distribution may be a consequence of the HST s attitude history and pointing requirements for the cold radiator, or of environmental effects, such as an anisotropic distribution of the responsible population in that size regime. Understanding the size-dependent spatial distribution of impact features is essential to the general analysis of these features. We have obtained from GSFC a 15 minute temporal resolution record of the state vector (Earth Centered Inertial position and velocity) and HST attitude, consisting of the orientation of the velocity and HST-sun vectors in HST body coordinates. This paper reviews the actual state vector and attitude history of the radiator in the context of the randomly tumbling plate assumption and assesses the statistical likelihood (or collection efficiency) of the radiator for the micrometeoroid and orbital debris environments. The NASA Marshall Space Flight Center s Meteoroid Environment Model is used to assess the micrometeoroid component. The NASA Orbital Debris Engineering Model (ORDEM) is used to model the orbital debris component. Modeling results are compared with observations of the impact feature spatial distribution, and the relative contribution of each environmental component are examined in detail.

Anz-Meador, Phillip D.↗

NASA NTRS2010

Progress in Development of the Axel Rovers

Progress has been made in the development of a family of robotic land vehicles having modular and minimalist design features chosen to impart a combination of robustness, reliability, and versatility. These vehicles at earlier stages of development were described in two previous NASA Tech Briefs articles: "Reconfigurable Exploratory Robotic Vehicles" (NPO-20944), Vol. 25, No. 7 (July 2001), page 56; and "More About Reconfigurable Exploratory Robotic Vehicles" (NPO-30890), Vol. 33, No. 8 (August 2009), page 40. Conceived for use in exploration of the surfaces of Mars and other remote planets, these vehicles could also be adapted to terrestrial applications, including exploration of volcanic craters or other hostile terrain, military reconnaissance, inspection of hazardous sites, and searching for victims of earthquakes, landslides, avalanches, or mining accidents. In addition, simplified versions of these vehicles might be marketable as toys. The most basic module in this family of reconfigurable robots is the Axel rover, which has a cylindrical body with two main wheels and a trailing link. Inside its body are three motors and associated mechanisms for driving the two wheels and for rotating the link 360 around its symmetrical body. The actuated link serves several purposes: It is used as a lever arm to react to the wheels thrust to move Axel in multiple directions. It is used to rotate the Axel housing in order to tilt, to the desired angle, any sensors and instruments mounted on or in the Axel housing. It provides an alternative mobility mode, which is primarily used in its tethered configuration. Turn ing the link into the ground in lieu of driving the wheels causes the Axel housing and wheels to roll as a unit and thereby leads to a tumbling motion along the ground. With a tether mounted around Axel s cylindrical body, the link serves as a winch mechanism to reel and unreel the tether raising and lowering Axel over steep and vertical surfaces (Figure 1). Sensors, computation, and communication modules are also housed inside Axel s body. A pair of stereo vision cameras provides three-dimensional view for autonomous navigation and avoiding obstacles. Inertial sensors determine the tilt of the robot and are used for estimating its motion. In a fully developed version, power would be supplied by rechargeable batteries aboard Axel; at the time of reporting the information for this article, power was supplied from an external source via a cable. In and of itself, the Axel rover is fully capable of traversing and sampling terrains on planetary surfaces. By use of only the two main wheel actuators and the caster link actuator, Axel can be made to follow an arbitrary path, turn in place, and operate upside- down or right-side-up. If operated in a tethered configuration, as shown in Figure 1, it can be made to move down and up a steep crater wall, descend from an overhang to a cave, and ascend from the cave back to the overhang, all by use of the same three actuators. Such tethered operation could be useful in searching for accident victims or missing persons in mines, caves, and rubble piles. Running the tether through the caster link enhances the stability of Axel and provides a restoring force that keeps the link off the ground for the most part during operation on a steep slope. In its extended configuration, two Axel modules can dock to either side of a payload module to form the four wheeled Axel2 rover (Figure 2). Additional payload and Axel modules can dock to either side of the Axel2 to form the Axel3 rover, extending its payload capacity and its mobility capabilities.

Nesnas, Issa A.↗

NASA NTRS2011

Engineering Challenges for Active Debris Removal

Recent modeling studies on the instability of the debris population in the low Earth orbit (LEO) region and the collision between Iridium 33 and Cosmos 2251 have underlined the need for active debris removal. A 2009 analysis by the NASA Orbital Debris Program Office shows that, in order to maintain the LEO debris population at a constant level for the next 200 years, an active debris removal of about five objects per year is needed. The targets identified for removal are those with the highest mass and collision probability products in the environment. Many of these objects are spent upper stages with masses ranging from 1 to more than 8 metric tons, residing in several altitude regions and concentrated in about 10 inclination bands. To remove five of those objects on a yearly basis, in a cost-effective manner, represents many challenges in engineering, technology development, and operations. This paper outlines a conceptual end-to-end debris removal operation, including launch, precision tracking, rendezvous, stabilization (of the tumbling targets), capture, and deorbit of the targets; and highlights major challenges associated with the operations. Pros and cons of several proposed removal techniques are also evaluated.

Liou, Jer-Chyi↗