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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 217 records · Page 12
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NASA NTRS1991

ELV payload environment; Proceedings of the NASA-Industry Conference on Launch Environments of ELV Payloads, Elkridge, MD, June 4-6, 1990

Various papers on the ELV payload environment are presented. Individual topics addressed include: EnviroNET: the Space Environment Information Service, ELV Payload Environments Study, perspective of GSFC Orbital Launch Services Office, Cosmic Background Explorer, the Standard Booster Interface, MLV Standard Interface Study, future payload requirements, General Dynamics Atlas family, Delta II Launch Vehicle environments, Titan vibroacoustics, Titan IV, the Scout Launch Vehicle, Pegasus.

Lauriente, Michael↗
NASA NTRS1991

Systems engineering studies of on-orbit assembly operation

While the practice of construction has a long history, the underlying theory of construction is relatively young. Very little has been documented as to techniques of logistic support, construction planning, construction scheduling, construction testing, and inspection. The lack of 'systems approaches' to construction processes is certainly one of the most serious roadblocks to the construction of space structures. System engineering research efforts at CSC are aimed at developing concepts and tools which contribute to a systems theory of space construction. The research is also aimed at providing means for trade-offs of design parameters for other research areas in CSC. Systems engineering activity at CSC has divided space construction into the areas of orbital assembly, lunar base construction, interplanetary transport vehicle construction, and Mars base construction. A brief summary of recent results is given. Several models for 'launch-on-time' were developed. Launch-on-time is a critical concept to the assembly of such Earth-orbiting structures as the Space Station Freedom, and to planetary orbiters such as the Mars transfer vehicle. CSC has developed a launch vehicle selection model which uses linear programming to find optimal combinations of launch vehicles of various sizes (Atlas, Titan, Shuttles, HLLV's) to support SEI missions. Recently, the Center developed a cost trade-off model for studying on orbit assembly logistics. With this model it was determined that the most effective size of the HLLV would be in the range of 120 to 200 metric tons to LEO, which is consistent with the choices of General Stafford's Synthesis Group Report. A second-generation Dynamic Construction Activities Model ('DYCAM') process model has been under development, based on our past results in interruptability and our initial DYCAM model. This second-generation model is built on the paradigm of knowledge-based expert systems. It is aimed at providing answers to two questions: (1) what are some necessary or sufficient conditions for judging conceptual designs of spacecraft?, and (2) can a methodology be formulated such that these conditions may be used to provide computer-aided tools for evaluating conceptual designs and planning for space assembly sequences? Early simulation results indicate that the DYCAM model has a clear ability to emulate and simulate human orbital construction processes.

Morgenthaler, George W.↗
NASA NTRS1993

Planetary Mission Capability of Small Low-Power Solar Electric Propulsion Systems

There is currently a vigorous investigation underway of low-cost planetary missions using small, inexpensive spacecraft. In order to keep the total mission costs down either a medium class launch vehicle such as a Delta II or an intermediate class launch vehicle such as an Atlas IIA or Atlas IIAS would be required for these planetary missions.

Low-Power Solar Electric↗
NASA NTRS2012

Returning an Entire Near-Earth Asteroid in Support of Human Exploration Beyond Low-Earth Orbit

This paper describes the results of a study into the feasibility of identifying, robotically capturing, and returning an entire Near-Earth Asteroid (NEA) to the vicinity of the Earth by the middle of the next decade. The feasibility of such an asteroid retrieval mission hinges on finding an overlap between the smallest NEAs that could be reasonably discovered and characterized and the largest NEAs that could be captured and transported in a reasonable flight time. This overlap appears to be centered on NEAs roughly 7 m in diameter corresponding to masses in the range of 250,000 kg to 1,000,000 kg. The study concluded that it would be possible to return a approx.500,000-kg NEA to high lunar orbit by around 2025. The feasibility is enabled by three key developments: the ability to discover and characterize an adequate number of sufficiently small near-Earth asteroids for capture and return; the ability to implement sufficiently powerful solar electric propulsion systems to enable transportation of the captured NEA; and the proposed human presence in cislunar space in the 2020s enabling exploration and exploitation of the returned NEA. Placing a 500-t asteroid in high lunar orbit would provide a unique, meaningful, and affordable destination for astronaut crews in the next decade. This disruptive capability would have a positive impact on a wide range of the nation's human space exploration interests. It would provide a high-value target in cislunar space that would require a human presence to take full advantage of this new resource. It would offer an affordable path to providing operational experience with astronauts working around and with a NEA that could feed forward to much longer duration human missions to larger NEAs in deep space. It represents a new synergy between robotic and human missions in which robotic spacecraft would retrieve significant quantities of valuable resources for exploitation by astronaut crews to enable human exploration farther out into the solar system. The capture, transportation, examination, and dissection of an entire NEA would provide valuable information for planetary defense activities that may someday have to deflect a much larger near-Earth object. Transportation of the NEA to lunar orbit with a total flight time of 6 to 10 years would be enabled by a ~40-kW solar electric propulsion system with a specific impulse of 3,000 s. The flight system could be launched to low-Earth orbit (LEO) on a single Atlas V-class launch vehicle, and return to lunar orbit a NEA with at least 28 times the mass launched to LEO. Longer flight times, higher power SEP systems, or a target asteroid in a particularly favorable orbit could increase the mass amplification factor from 28-to-1 to 70-to-1 or greater. The NASA GRC COMPASS team estimated the full life-cycle cost of an asteroid capture and return mission at approx.$2.6B.

solar electric propulsion↗
NASA NTRS2010

Statistical Analysis of Model Data for Operational Space Launch Weather Support at Kennedy Space Center and Cape Canaveral Air Force Station

The 12-km resolution North American Mesoscale (NAM) model (MesoNAM) is used by the 45th Weather Squadron (45 WS) Launch Weather Officers at Kennedy Space Center (KSC) and Cape Canaveral Air Force Station (CCAFS) to support space launch weather operations. The 45 WS tasked the Applied Meteorology Unit to conduct an objective statistics-based analysis of MesoNAM output compared to wind tower mesonet observations and then develop a an operational tool to display the results. The National Centers for Environmental Prediction began running the current version of the MesoNAM in mid-August 2006. The period of record for the dataset was 1 September 2006 - 31 January 2010. The AMU evaluated MesoNAM hourly forecasts from 0 to 84 hours based on model initialization times of 00, 06, 12 and 18 UTC. The MesoNAM forecast winds, temperature and dew point were compared to the observed values of these parameters from the sensors in the KSC/CCAFS wind tower network. The data sets were stratified by model initialization time, month and onshore/offshore flow for each wind tower. Statistics computed included bias (mean difference), standard deviation of the bias, root mean square error (RMSE) and a hypothesis test for bias = O. Twelve wind towers located in close proximity to key launch complexes were used for the statistical analysis with the sensors on the towers positioned at varying heights to include 6 ft, 30 ft, 54 ft, 60 ft, 90 ft, 162 ft, 204 ft and 230 ft depending on the launch vehicle and associated weather launch commit criteria being evaluated. These twelve wind towers support activities for the Space Shuttle (launch and landing), Delta IV, Atlas V and Falcon 9 launch vehicles. For all twelve towers, the results indicate a diurnal signal in the bias of temperature (T) and weaker but discernable diurnal signal in the bias of dewpoint temperature (T(sub d)) in the MesoNAM forecasts. Also, the standard deviation of the bias and RMSE of T, T(sub d), wind speed and wind direction indicated the model error increased with the forecast period all four parameters. The hypothesis testing uses statistics to determine the probability that a given hypothesis is true. The goal of using the hypothesis test was to determine if the model bias of any of the parameters assessed throughout the model forecast period was statistically zero. For th is dataset, if this test produced a value >= -1 .96 or <= 1.96 for a data point, then the bias at that point was effectively zero and the model forecast for that point was considered to have no error. A graphical user interface (GUI) was developed so the 45 WS would have an operational tool at their disposal that would be easy to navigate among the multiple stratifications of information to include tower locations, month, model initialization times, sensor heights and onshore/offshore flow. The AMU developed the GUI using HyperText Markup Language (HTML) so the tool could be used in most popular web browsers with computers running different operating systems such as Microsoft Windows and Linux.

Bauman, William H., III↗
NASA NTRS1992

Design definition of the Laser Atmospheric Wind Sounder (LAWS), phase 2. Volume 1: Executive summary

The LAWS phase 1 and phase 2 studies have been completed on schedule and have led to significant advances in CO2 laser development. The Phase 2 Design Definition Study has shown that a large scanning mirror/high pulse energy laser LAWS Instrument is feasible and within the existing technology. The capability to monitor wind velocities with backscatter ratios of 10(exp 11) m(exp -1) SR(exp -1) is feasible. The weight budget allocated for the baseline LAWS is adequate, and sufficient reserves exist with the potential downsized configuration. With the possible decrease in available power from the baseline of 2.2 kW guideline, power and shot management is critical for the baseline configuration (15 to 20 J). This is particularly true during the 100 day occultation period each year. With the downsized configurations (5 to 7 J), power management is still necessary during the occultation but is primarily limited to shot management over the polar regions. The breadboard effort has produced significant laser advances for a tight 18 month schedule and the minimum budgets available from NASA, Lockheed, and TDS. Using the NASA funds and Lockheed and TDS fixed assets budgets, the breadboard was designed, fabricated, and brought on-line with first laser light within 16 months after ATP. First laser beam was obtained on 21 April 1992 at a 5 J power level. Tests since then have been conducted at sustained, repetitive pulse levels of over 7 J and 20 Hz. This is an increase of over two or three times greater than any system previously developed from this type laser. Increased power levels and additional life tests will be accomplished in the next LAWS phase. The Lockheed LAWS design will operate in the gravity gradient mode on-orbit, and all possible instrument vibration and jitter modes have been considered. Adequate pointing stability and control is state-of-the-art technology for the critical time periods, frequency rates, and control responses required by LAWS. Lockheed recommends a 6-1/2 year phase C/D program for LAWS to provide adequate feedback from the engineering unit and the qualification unit to the final flight unit. Assuming a one year period for LAWS integration to the spacecraft, followed by a six-month period for launch vehicle integration, LAWS could be successfully developed and launched in eight years. Our baseline design or downsized design can be accommodated by either the Atlas 2AS or the Delta launch vehicles. Lockheed's recommendation is that, based on the successful phase 2 design study and breadboard program, a follow-on 18 month extended breadboard testing program and additional system engineering studies, primarily in interfacing with a to be defined platform, be initiated. This should be immediately followed by the phase C/D program, leading to a LAWS launch in late 2001 or early 2002.

Source record↗
NASA NTRSDate not supplied

Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) Aeroshell Performance

On November 10, 2022, NASA launched the Low-earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) reentry vehicle (RV)as a secondary payload mounted inside the launch vehicle (LV) adaptor on an Atlas V 401 out of the Vandenburg Space Force Base (VSFB). The primary payload, Joint Polar Satellite System-2 (JPSS-2), was delivered successfully to a sun-synchronous trajectory shortly after launch, at which point the Centaur upper stage performed a burn to de-orbit the system. Once on the desired trajectory to enter the atmosphere, the top of the payload adaptor was ejected to expose the LOFTID RV. Next, the LOFTID aeroshell, a Hypersonic Inflatable Aerodynamic Decelerator (HIAD), was deployed and inflated without issue. The Centaur pointed the LOFTID RV to the desired attitude to enter the atmosphere, spun the assembly to three rpm, separated the LOFTID RV, and finally performed a divert maneuver to prevent the Centaur from re-contacting the LOFTID RV after atmospheric entry. The LOFTID RV maintained the design attitude and spin rate from separation to atmospheric entry, a duration of roughly 30 minutes, demonstrating the aeroshell did not experience any energy damping from inflatable structure flexibility. The LOFTID RV entered the atmosphere over Alaska at >8km/sec and decelerated as designed, demonstrating aerodynamically stable flight from hypersonic entry through subsonic parachute deployment. On-board visible-spectrum cameras captured the physical response of the heatshield through all phases of flight. Co-located infrared cameras captured the temperature distribution of the aftside of the heat shield anchored to a distribution of thermocouples on the inflatable structure (IS) in the field of view. Thermocouples were also embedded in the forward side of the aeroshell both in the Flexible Thermal Protection System (FTPS) as well as the IS. This paper will document aeroshell’s response to the atmospheric entry. The visible-spectrum cameras captured mechanical response of the deployable aeroshell to the loads encountered in the different phases of entry. Events seen on the visible-spectrum cameras correlate with the measurements of the load cell pins on webbing elements that attach the aeroshell to the centerbody structure. Thermocouple data captured verifies the aerothermal response of the aeroshell was in-kind with pre-flight analysis predictions although somewhat lower in most locations. This data will enable enhancement of HIAD predictive capability. Post-flight inspection of the aeroshell after splashdown and ocean recovery shows that all the aeroshell materials and manufacturing techniques developed over the past two decades of HIAD technology development performed as expected at large scale, a scale which cannot be replicated in ground test facilities. The LOFTID RV remained high in the water after touching down under parachute, and maintained buoyancy and integrity for the hour it took to position the recovery vessel near the vehicle after splashdown. There was no indication the HIAD would have experienced an issue maintaining buoyancy for many more hours after water impact as the component maintained adequate inflation pressure more than 24 hours after water extraction with no additional inflation gas in tanks, as the inflation gas remaining in the tank was purposely vented prior to splashdown.

HIAD↗
NASA NTRS2011

Space Transportation Infrastructure Supported By Propellant Depots

A space transportation infrastructure is described that utilizes propellant depots to support all foreseeable missions in the Earth-Moon vicinity and deep space out to Mars. The infrastructure utilizes current expendable launch vehicles such as the Delta IV Heavy, Atlas V, and Falcon 9, for all crew, cargo, and propellant launches to orbit. Propellant launches are made to a Low-Earth-Orbit (LEO) Depot and an Earth-Moon Lagrange Point 1 (L1) Depot to support new reusable in-space transportation vehicles. The LEO Depot supports missions to Geosynchronous Earth Orbit (GEO) for satellite servicing, and to L1 for L1 Depot missions. The L1 Depot supports Lunar, Earth-Sun L2 (ESL2), Asteroid, and Mars missions. A Mars Orbital Depot is also described to support ongoing Mars missions. New concepts for vehicle designs are presented that can be launched on current 5-meter diameter expendable launch vehicles. These new reusable vehicle concepts include a LEO Depot, L1 Depot, and Mars Orbital Depot based on International Space Station (ISS) heritage hardware. The high-energy depots at L1 and Mars orbit are compatible with, but do not require, electric propulsion tug use for propellant and/or cargo delivery. New reusable in-space crew transportation vehicles include a Crew Transfer Vehicle (CTV) for crew transportation between the LEO Depot and the L1 Depot, a new reusable Lunar Lander for crew transportation between the L1 Depot and the lunar surface, and a Deep Space Habitat (DSH) to support crew missions from the L1 Depot to ESL2, Asteroid, and Mars destinations. A 6 meter diameter Mars lander concept is presented that can be launched without a fairing based on the Delta IV heavy Payload Planners Guide, which indicates feasibility of a 6.5 meter fairing. This lander would evolve to re-usable operations when propellant production is established on Mars. Figure 1 provides a summary of the possible missions this infrastructure can support. Summary mission profiles are presented for each primary mission capability. These profiles are the basis for propellant loads, numbers of vehicles/stages and launches for each mission capability. Data includes the number of launches required for each mission utilizing current expendable launch vehicle systems, and concluding remarks include ideas for reducing the number of launches through incorporation of heavy-lift launch vehicles, solar electric propulsion, and other transportation support concepts.

Smitherman, David↗
NASA NTRS2012

ELaNa - Educational Launch of Nanosatellite Providing Routine RideShare Opportunities

Since the creation of the NASA CubeSat Launch Initiative (NCSLI), the need for CubeSat rideshares has dramatically increased. After only three releases of the initiative, a total of 66 CubeSats now await launch opportunities. So, how is this challenge being resolved? NASA's Launch Services Program (LSP) has studied how to integrate PPODs on Athena, Atlas V, and Delta IV launch vehicles and has been instrumental in developing several carrier systems to support CubeSats as rideshares on NASA missions. In support of the first two ELaNa missions the Poly-Picosatellite Orbital Deployer (P-POD) was adapted for use on a Taurus XL (ELaNa I) and a Delta n (ELaNa III). Four P-PODs, which contained a total eight CubeSats, were used on these first ELaNa missions. Next up is ELaNa VI, which will launch on an Atlas V in August 2012. The four ELaNa VI CubeSats, in three P-PODs, are awaiting launch, having been integrated in the NPSCuLite. To increase rideshare capabilities, the Launch Services Program (LSP) is working to integrate P-PODs on Falcon 9 missions. The proposed Falcon 9 manifest will provide greater opportunities for the CubeSat community. For years, the standard CubeSat size was 1 U to 3U. As the desire to include more science in each cube grows, so does the standard CubeSat size. No longer is a 1 U, 1.5U, 2U or 3U CubeSat the only option available; the new CubeSat standard will include 6U and possibly even 12U. With each increase in CubeSat size, the CubeSat community is pushing the capability of the current P-POD design. Not only is the carrier system affected, but integration to the Launch Vehicle is also a concern. The development of a system to accommodate not only the 3U P-POD but also carriers for larger CubeSats is ongoing. LSP considers payloads in the lkg to 180 kg range rideshare or small/secondary payloads. As new and emerging small payloads are developed, rideshare opportunities and carrier systems need to be identified and secured. The development of a rideshare carrier system is not always cost effective. Sometimes a launch vehicle with an excellent performance record appears to be a great rideshare candidate however, after completing a feasibility study, LSP may determine that the cost of the rideshare carrier system is too great and, due to budget constraints, the development cannot go forward. With the current budget environment, one cost effective way to secure rideshare opportunities is to look for synergy with other government organizations that share the same interest.

Skrobot, Garrett Lee↗
NASA NTRS2015

RL10 Engine Ability to Transition from Atlas to Shuttle/Centaur Program

A key launch vehicle design feature is the ability to take advantage of new technologies while minimizing expensive and time consuming development and test programs. With successful space launch experiences and the unique features of both the National Aeronautics and Space Administration (NASA) Space Transportation System (Space Shuttle) and Atlas/Centaur programs, it became attractive to leverage these capabilities. The Shuttle/Centaur Program was created to transition the existing Centaur vehicle to be launched from the Space Shuttle cargo bay. This provided the ability to launch heaver and larger payloads, and take advantage of new unique launch operational capabilities. A successful Shuttle/Centaur Program required the Centaur main propulsion system to quickly accommodate the new operating conditions for two new Shuttle/Centaur configurations and evolve to function in the human Space Shuttle environment. This paper describes the transition of the Atlas/Centaur RL10 engine to the Shuttle/Centaur configurations; shows the unique versatility and capability of the engine; and highlights the importance of ground testing. Propulsion testing outcomes emphasize the value added benefits of testing heritage hardware and the significant impact to existing and future programs.

Launch Vehicle↗
NASA NTRS1999

Advanced Electric Propulsion for RLV Launched Geosynchronous Spacecraft

Solar Electric Propulsion (SEP) when used for station keeping and final orbit insertion has been shown to increase a geostationary satellite's payload when launched by existing expendable launch vehicles. In the case of reusable launch vehicles or expendable launch vehicles where an upper stage is an expensive option, this methodology can be modified by using the existing on-board apogee chemical system to perform a perigee burn and then letting the electric propulsion system complete the transfer to geostationary orbit. The elimination of upper stages using on-board chemical and electric propulsion systems was thus examined for GEO spacecraft. Launch vehicle step-down from an Atlas IIAR to a Delta 7920 (no upper stage) was achieved using expanded on-board chemical tanks, 40 kW payload power for electric propulsion, and a 60 day elliptical to GEO SEP orbit insertion. Optimal combined chemical and electric trajectories were found using SEPSPOT. While Hall and ion thrusters provided launch vehicle step-down and even more payload for longer insertion times, NH3 arcjets had insufficient performance to allow launch vehicle step-down. Degradation levels were only 5% to 7% for launch step-down cases using advanced solar arrays. Results were parameterized to allow comparisons for future reusable launch vehicles. Results showed that for an 8 W/kg initial power/launch mass power density spacecraft, 50% to 100% more payload can be launched using this method.

Oleson, Steven↗
NASA NTRS1988

The Atlas-Centaur 67 incident

The conditions leading to the breakup of the Atlas-Centaur 67 (AC-67) vehicle launched on March 26, 1987 during a typical winter-time storm are analyzed, and a most probable lightning strike scenario was developed based on inspection of the AC-67 debris, the AC-68 vehicle, and appropriate drawings, electrical diagrams, and photographs. It is shown that, during ascent, the vehicle encountered increasingly larger electric fields and flew through clouds that produced precipitation static on the vehicle, eventually triggering a cloud-to-ground lightning, comprised of at least four return strokes. The resulting lightning current coupled a signal into the wiring which goes to the AC-67 digital computer unit (DCU), effecting a single-word memory alteration and causing the DCU to issue a hardover engine gimbal command. This led to an excessive angle of attack, large dynamic loads, and the breakup of the AC-67.

Christian, H. J.↗
NASA NTRS1992

An electromechanical actuation system for an expendable launch vehicle

A major effort at the NASA Lewis Research Center in recent years has been to develop electro-mechanical actuators (EMA's) to replace the hydraulic systems used for thrust vector control (TVC) on launch vehicles. This is an attempt ot overcome the inherent inefficiencies and costs associated with the existing hydraulic structures. General Dynamics Space Systems Division, under contract to NASA Lewis, is developing 18.6 kW (25 hp), 29.8 kW (40 hp), and 52.2 kW (70 hp) peak EMA systems to meet the power demands for TVC on a family of vehicles developed for the National Launch System. These systems utilize a pulse population modulated converter and field-oriented control scheme to obtain independent control of both the voltage and frequency. These techniques allow an induction motor to be operated at its maximum torque at all times. At NASA Lewis, we are building on this technology to develop our own in-house system capable of meeting the peak power requirements for an expendable launch vehicle (ELV) such as the Atlas. Our EMA will be capable of delivering 22.4 kW (30 hp) peak power with a nominal of 6.0 kW (8 hp). This system differs from the previous ones in two areas: (1) the use of advanced control methods, and (2) the incorporation of built-in-test. The advanced controls are essential for minimizing the controller size, while the built-in-test is necessary to enhance the system reliability and vehicle health monitoring. The ultimate goal of this program is to demonstrate an EMA which will be capable of self-test and easy integration into other projects. This paper will describe the effort underway at NASA Lewis to develop an EMA for an Atlas class ELV. An explanation will be given for each major technology block, and the status of each major technology block and the status of the overall program will be reported.

Burrows, Linda M.↗
NASA NTRS2009

The International X-ray Observatory

The International X-Ray Observatory, a joint NASA-ESA-JAXA effort, is a next generation X-ray telescope that will answer many fundamental questions in contemporary astrophysics such as how do supermassive black holes influence galaxy evolution and how do galaxy clusters evolve (and how does this constrain dark energy and dark matter)? As a powerful astronomical observatory, IXO will also address questions ranging from the neutron star equation of state to the distribution and dynamical state of intergalactic material. X-ray spectroscopy, polarimetry, and timing studies provided by IXO's instruments will give detailed measures of abundances, temperatures, densities, magnetic fields and gravitational potentials. These measurements will be complementary to the next generation of observatories such as ALMA, JWST, and future ground-based optical-NIR telescopes. This mission will be ready for launch in the 2020-2021 timeframe and will launch on an Atlas V or Ariane V launch vehicle to L2. It employs a deployable optical bench to achieve the 20 meter focal length and a suite of five instruments. This talk will describe the motivating science for this mission as well as the spacecraft, instruments and optics

Hornschemeier, A.↗
NASA NTRS2013

2011 Mars Science Laboratory Trajectory Reconstruction and Performance from Launch Through Landing

The Mars Science Laboratory (MSL) mission successfully launched on an Atlas V 541 Expendable Evolved Launch Vehicle (EELV) from the Eastern Test Range (ETR) at Cape Canaveral Air Force Station (CCAFS) in Florida at 15:02:00 UTC on November 26th, 2011. At 15:52:06 UTC, six minutes after the MSL spacecraft separated from the Centaur upper stage, the spacecraft transmitter was turned on and in less than 20 s spacecraft carrier lock was achieved at the Universal Space Network (USN) Dongara tracking station located in Western Australia. MSL, carrying the most sophisticated rover ever sent to Mars, entered the Martian atmosphere at 05:10:46 SpaceCraft Event Time (SCET) UTC, and landed inside Gale Crater at 05:17:57 SCET UTC on August 6th, 2012. Confirmation of nominal landing was received at the Deep Space Network (DSN) Canberra tracking station via the Mars Odyssey relay spacecraft at 05:31:45 Earth Received Time (ERT) UTC. This paper summarizes in detail the actual vs. predicted trajectory performance in terms of launch vehicle events, launch vehicle injection performance, actual DSN/USN spacecraft lockup, trajectory correction maneuver performance, Entry, Descent, and Landing events, and overall trajectory and geometry characteristics.

trajectory↗
NASA NTRS2021

Mars 2020 Perseverance trajectory reconstruction and performance from launch through landing

The Mars 2020 (M2020) Mission carrying Perseverance, the most advanced rover ever sent to Mars, successfully launched on an Atlas V 541 (AV-088) launch vehicle from the Eastern Test Range (ETR) at Cape Canaveral Air Force Station (CCAFS) in Florida at 11:50:00 UTC (T-Zero time) on July 30, 2020. After some station reconfiguration, carrier/telemetry were locked at both Deep Space Network (DSN) Canberra and Goldstone stations. Perseverance entered the Martian atmosphere at 20:36:50 Spacecraft Event Time (SCET) UTC, and landed inside Jezero Crater at 20:43:49 SCET UTC on February 18, 2021. Confirmation of nominal landing was received at the DSN Goldstone and Madrid tracking stations via the Mars Reconnaissance Orbiter at 20:55:11 Earth Received Time (ERT) UTC. This paper summarizes in detail the actual vs. predicted performance in terms of launch vehicle events, launch vehicle injection performance, actual DSN spacecraft lockup, trajectory correction maneuver performance, Entry, Descent, and Landing events, and overall trajectory and geometric characteristics.

Abilleira, Fernando↗
NASA NTRS2021

Mars 2020 Perseverance Trajectory Reconstruction and Performance from Launch through Landing

The Mars 2020 (M2020) Mission carrying Perseverance, the most advanced rover ever sent to Mars, successfully launched on an Atlas V 541 (AV-088) launch vehicle from the Eastern Test Range (ETR) at Cape Canaveral Air Force Station (CCAFS) in Florida at 11:50:00 UTC (T-Zero time) on July 30, 2020. After some station reconfiguration, carrier/telemetry were locked at both Deep Space Network (DSN) Canberra and Goldstone stations. Perseverance entered the Martian atmosphere at 20:36:50 Spacecraft Event Time (SCET) UTC, and landed inside Jezero Crater at 20:43:49 SCET UTC on February 18, 2021. Confirmation of nominal landing was received at the DSN Goldstone and Madrid tracking stations via the Mars Reconnaissance Orbiter at 20:55:11 Earth Received Time (ERT) UTC. This paper summarizes in detail the actual vs. predicted performance in terms of launch vehicle events, launch vehicle injection performance, actual DSN spacecraft lockup, trajectory correction maneuver performance, Entry, Descent, and Landing events, and overall trajectory and geometric characteristics.

Wong, Mau↗
NASA NTRS2012

FASTSAT a Mini-Satellite Mission...A Way Ahead

The Fast Affordable Science and Technology Spacecraft (FASTSAT) is a mini-satellite weighing less than 150 kg. FASTSAT was developed as government-industry collaborative research and development flight project targeting rapid access to space to provide an alternative, low cost platform for a variety of scientific, research, and technology payloads. The initial spacecraft was designed to carry six instruments and launch as a secondary rideshare payload. This design approach greatly reduced overall mission costs while maximizing the on-board payload accommodations. FASTSAT was designed from the ground up to meet a challenging short schedule using modular components with a flexible, configurable layout to enable a broad range of payloads at a lower cost and shorter timeline than scaling down a more complex spacecraft. The integrated spacecraft along with its payloads were readied for launch 15 months from authority to proceed. As an ESPA-class spacecraft, FASTSAT is compatible with many different launch vehicles, including Minotaur I, Minotaur IV, Delta IV, Atlas V, Pegasus, Falcon 1/1e, and Falcon 9. These vehicles offer an array of options for launch sites and provide for a variety of rideshare possibilities.

Boudreaux, Mark↗