Search NASA⌕ Search

SEARCH · Search NASA

Results for “Mission Elapsed Time”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 37 records · Page 2

Anomaly Trends for Missions to Mars: Mars Global Surveyor and Mars Odyssey

The long term flight operations of the Mars Global Surveyor and Mars Odyssey spacecraft give us an excellent chance to examine the operations of two long lived spacecraft in orbit around Mars during overlapping time periods. This study examined the anomalies for each mission maintained for NASA at the Jet Propulsion Laboratory. By examining the anomalies each mission encountered during their multiyear missions, trends were identified related to when anomalies occurred during each mission, the types of anomalies encountered, and corrective actions taken to mitigate the effects of the anomalies. As has been discovered in previous studies the numbers of anomalies directly correlate with mission activity and show a decreasing trend with elapsed mission time. Trend analysis also identified a heavy emphasis on software as the source or solution to anomalies for both missions.

Mars↗

Three Orbital Burns to Molniya Orbit via Shuttle Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982-86 joint NASA-USAF Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57deg inclined orbit: 9,545 lb vs. 9,552 lb of separated spacecraft weight respectively. There was a significant reduction in the need for propellant launch time reserve for a one hour window: only 78 lb for the three burn transfer vs. 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three vs. two burn transfer (12 vs. 11/4 hrs), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗

Three Orbital Burns to Molniya Orbit Via Shuttle_Centaur G Upper Stage

An unclassified analytical trajectory design, performance, and mission study was done for the 1982 to 1986 joint National Aeronautics and Space Administration (NASA)-United States Air Force (USAF) Shuttle/Centaur G upper stage development program to send performance-demanding payloads to high orbits such as Molniya using an unconventional orbit transfer. This optimized three orbital burn transfer to Molniya orbit was compared to the then-baselined two burn transfer. The results of the three dimensional trajectory optimization performed include powered phase steering data and coast phase orbital element data. Time derivatives of the orbital elements as functions of thrust components were evaluated and used to explain the optimization's solution. Vehicle performance as a function of parking orbit inclination was given. Performance and orbital element data was provided for launch windows as functions of launch time. Ground track data was given for all burns and coasts including variation within the launch window. It was found that a Centaur with fully loaded propellant tanks could be flown from a 37 deg inclination low Earth parking orbit and achieve Molniya orbit with comparable performance to the baselined transfer which started from a 57 deg inclined orbit: 9,545 versus 9,552 lb of separated spacecraft weight, respectively. There was a significant reduction in the need for propellant launch time reserve for a 1 hr window: only 78 lb for the three burn transfer versus 320 lb for the two burn transfer. Conversely, this also meant that longer launch windows over more orbital revolutions could be done for the same amount of propellant reserve. There was no practical difference in ground tracking station or airborne assets needed to secure telemetric data, even though the geometric locations of the burns varied considerably. There was a significant adverse increase in total mission elapsed time for the three versus two burn transfer (12 vs. 1-1/4 hr), but could be accommodated by modest modifications to Centaur systems. Future applications were discussed. The three burn transfer was found to be a viable, arguably preferable, alternative to the two burn transfer.

Orbital trajectory design↗

Ballistic Leverage for Conjunction Avoidance During the Lunar Transit Trajectory of NASA's Co-Manifested Vehicle

Analysis is completed to investigate the sensitivity of the Lunar Transit Trajectory of NASA’s Co-Manifested Vehicle (CMV) to the insertion of intentional coasting into an otherwise nearly continuous low thrust maneuver. A reference trajectory is developed which explicitly models low altitude thruster shutdowns in a level of detail not included in previous Design Reference Missions. This reference trajectory is used to investigate the sensitivity of the Lunar Transit to additional inserted coasting, presented in the form of separation of the perturbed trajectory from the reference as a function of mission elapsed time and duration and location of the coast segment. These results are shown to inform potential conjunction avoidance strategies during the CMV’s long low thrust spiral from a highly elliptic Earth parking orbit to the target Gateway Near Rectilinear HALO Orbit (NRHO).

low thrust↗

Ballistic Leverage for Conjunction Avoidance During the Lunar Transit Trajectory of NASA's Co-Manifested Vehicle

Analysis is completed to investigate the sensitivity of the Lunar Transit Trajectory of NASA’s Co-Manifested Vehicle (CMV) to the insertion of intentional coasting into an otherwise nearly continuous low thrust maneuver. A reference trajectory is developed which explicitly models low altitude thruster shutdowns in a level of detail not included in previous Design Reference Missions. This reference trajectory is used to investigate the sensitivity of the Lunar Transit to additional inserted coasting, presented in the form of separation of the perturbed trajectory from the reference as a function of mission elapsed time and duration and location of the coast segment. These results are shown to inform potential conjunction avoidance strategies during the CMV’s long low thrust spiral from a highly elliptic Earth parking orbit to the target Gateway Near Rectilinear HALO Orbit (NRHO).

low thrust↗

Modeling and Simulation of Variable Mass, Flexible Structures

The advent of the new Ares I launch vehicle has highlighted the need for advanced dynamic analysis tools for variable mass, flexible structures. This system is composed of interconnected flexible stages or components undergoing rapid mass depletion through the consumption of solid or liquid propellant. In addition to large rigid body configuration changes, the system simultaneously experiences elastic deformations. In most applications, the elastic deformations are compatible with linear strain-displacement relationships and are typically modeled using the assumed modes technique. The deformation of the system is approximated through the linear combination of the products of spatial shape functions and generalized time coordinates. Spatial shape functions are traditionally composed of normal mode shapes of the system or even constraint modes and static deformations derived from finite element models of the system. Equations of motion for systems undergoing coupled large rigid body motion and elastic deformation have previously been derived through a number of techniques [1]. However, in these derivations, the mode shapes or spatial shape functions of the system components were considered constant. But with the Ares I vehicle, the structural characteristics of the system are changing with the mass of the system. Previous approaches to solving this problem involve periodic updates to the spatial shape functions or interpolation between shape functions based on system mass or elapsed mission time. These solutions often introduce misleading or even unstable numerical transients into the system. Plus, interpolation on a shape function is not intuitive. This paper presents an approach in which the shape functions are held constant and operate on the changing mass and stiffness matrices of the vehicle components. Each vehicle stage or component finite element model is broken into dry structure and propellant models. A library of propellant models is used to describe the distribution of mass in the fuel tank or Solid Rocket Booster (SRB) case for various propellant levels. Based on the mass consumed by the liquid engine or SRB, the appropriate propellant model is coupled with the dry structure model for the stage. Then using vehicle configuration data, the integrated vehicle model is assembled and operated on by the constant system shape functions. The system mode shapes and frequencies can then be computed from the resulting generalized mass and stiffness matrices for that mass configuration. The rigid body mass properties of the vehicle are derived from the integrated vehicle model. The coupling terms between the vehicle rigid body motion and elastic deformation are also updated from the constant system shape functions and the integrated vehicle model. This approach was first used to analyze variable mass spinning beams and then prototyped into a generic dynamics simulation engine. The resulting code was tested against Crew Launch Vehicle (CLV-)class problems worked in the TREETOPS simulation package and by Wilson [2]. The Ares I System Integration Laboratory (SIL) is currently being developed at the Marshall Space Flight Center (MSFC) to test vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment and certify that the integrated system is prepared for flight. The Ares I SIL utilizes the Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) tool to simulate the launch vehicle and stimulate avionics hardware. Due to the presence of vehicle control system filters and the thrust oscillation suppression system, which are tuned to the structural characteristics of the vehicle, ARTEMIS must incorporate accurate structural models of the Ares I launch vehicle. The ARTEMIS core dynamics simulation models the highly coupled nature of the vehicle flexible body dynamics, propellant slosh, and vehicle nozzle inertia effects combined with mass and flexible body properties that vary significant with time during the flight. All forces that act on the vehicle during flight must be simulated, including deflected engine thrust force, spatially distributed aerodynamic forces, gravity, and reaction control jet thrust forces. These forces are used to excite an integrated flexible vehicle, slosh, and nozzle dynamics model for the vehicle stack that simulates large rigid body translations and rotations along with small elastic deformations. Highly effective matrix math operations on a distributed, threaded high-performance simulation node allow ARTEMIS to retain up to 30 modes of flex for real-time simulation. Stage elements that separate from the stack during flight are propagated as independent rigid six degrees of freedom (6DOF) bodies. This paper will present the formulation of the resulting equations of motion, solutions to example problems, and describe the resulting dynamics simulation engine within ARTEMIS.

Tobbe, Patrick A.↗

Shuttle STS-2 mission communication systems RF coverage and performance predictions. Volume 1: Ascent

The RF communications capabilities and nominally expected performance for the ascent phase of the second orbital flight of the shuttle are provided. Predicted performance is given mainly in the form of plots of signal strength versus elapsed mission time for the STDN (downlink) and shuttle orbiter (uplink) receivers for the S-band PM and FM, and UHF systems. Performance of the NAV and landing RF systems is treated for RTLS abort, since in this case the spacecraft will loop around and return to the launch site. NAV and landing RF systems include TACAN, MSBLS, and C-band altimeter. Signal strength plots were produced by a computer program which combines the spacecraft trajectory, antenna patterns, transmit and receive performance characteristics, and system mathematical models. When available, measured spacecraft parameters were used in the predictions; otherwise, specified values were used. Specified ground station parameter values were also used. Thresholds and other criteria on the graphs are explained.

Porter, J. A.↗

Camera/photometer results

Camera/photometer systems have been constructed for analyzing particulate contamination on STS-2. The systems contained two 16-mm photographic cameras to make stereoscopic observation of contaminant particles and the background. During the STS-2 mission, 102 photographic exposures were obtained under conditions suitable for recording particulate contamination. Attention is given to a breakdown of STS-2 camera/photometer data with the number of particles observed per frame, a correlation of observed contamination with mission elapsed time broken into 5-hr segments, and a correlation of observed contamination with onboard spacecraft events.

Clifton, K. S.↗

SKIRT Space Shuttle glow experiment

This paper describes a spectrometer/radiometer experiment to obtain infrared, visible, and ultraviolet measurements of Space Shuttle glow. The payload, Spacecraft Kinetic Infrared Test (SKIRT), is a cryogenic circular variable filter infrared spectrometer with a number of infrared, visible, and ultraviolet radiometers covering the spectral range of 0.2-5.4 microns and 9.9-10.3 microns. It will measure Shuttle glow as a function of mission elapsed time, orbiter attitude, temperature, and orbiter events such as thruster firings. The measured data should have sufficient spectral resolution and sensitivity to identify molecular species contributing to Shuttle glow emissions. SKIRT is manifested on STS-39.

Ahmadjian, M.↗

On the effect of grain burnback on STS-SRM fragment velocity

Concerns raised during the Ulysses Final Safety Analysis Review (FSAR) process called the solid rocket motor (SRM) fragment velocity prediction model into question. The specific area of concern was that there was a section of the SRM casing which was exposed to SRM chamber pressure as the grain (fuel) was consumed. These questions centered on the velocity of fragments which originated from the field joint region given that failure occurred between 37 and 72 seconds mission elapsed time (MET). Two dimensional coupled Eulerian-Lagrangian calculations were performed to assess the hot gas flow field which resulted from SRM casing fragmentation. The fragment to gas interface-pressure time-history obtained from these analyses was reduced to a boundary condition algorithm which was applied to an explicit-time-integration, finite element, three dimensional shell model of the SRM casing and unburned fuel. The results of these calculations showed that the velocity of fragments originating in the field joint was adequately described by the range of velocities given in the Shuttle Data Book (1988). Based on these results, no further analyses were required, and approval was obtained from the Launch Abort Subpanel of the Interagency Nuclear Safety Review Panel to use the SRM fragment velocity environments presented in the Ulysses FSAR (1990).

Eck, Marshall B.↗

STS-80 Flight Day 2

On this second day of the STS-80 mission, the flight crew, Cmdr. Kenneth D. Cockrell, Pilot Kent V. Rominger, Mission Specialists, Tamara E. Jernigan, Thomas D. Jones, and F. Story Musgrave, complete the first major objective of the mission with the deployment of the Orbiting Retrievable Far and Extreme Ultraviolet Spectrometer (ORFEUS) on the reusable Shuttle Pallet Satellite. Release of ORFEUS from Columbia's robot arm came at 8 hours 15 minutes mission elapsed time. Three hours after the release, ground controllers inform the crew that the instrument package appears to be working properly. This begins two weeks of gathering data on the origin and makeup of stars.

Source record↗

Magellan Post Launch Mission Operation Report

Magellan was successfully launched by the Space Shuttle Atlantis from the Kennedy Space Center at 2:47 p.m. EDT on May 4, 1989. The Inertial Upper Stage (IUS) booster and attached Magellan Spacecraft were successfully deployed from Atlantis on Rev. 5 as planned, at 06:14 hrs Mission Elapsed Time (MET). The two IUS propulsion burns which began at 07:14 hrs MET and were completed at 07:22 hrs MET, placed the Magellan Spacecraft almost perfectly on its preplanned trajectory to Venus. The IUS was jettisoned at 07:40 hrs MET and Magellan telemetry was immediately acquired by the Deep Space Network (DSN). A spacecraft trajectory correction maneuver was performed on May 21 and the spacecraft is in the planned standard cruise configuration with all systems operating nominally. An initial attempt was made to launch Atlantis on April 28, 1989, but the launch was scrubbed at T-31 sec due to a failure of the liquid hydrogen recirculation pump on Space Shuttle Main Engine #1. The countdown had proceeded smoothly until T-20 min when the Magellan radio receiver "locked-on" the MIL 71 Unified S-Band (USB) transmission as the transmitter power was increased fro 2 kw to 10 kw in support of the orbiter launch. During the planned hold at T-9 min, the USB was confirmed as the source of the receiver "lock" and Magellan's launch readiness was reaffirmed. In addition a five-minute extension of the T-9 hold occurred when a range safety computer went off-line, creating a loss of redundancy in the range safety computer network. Following resumption of the countdown, both the orbiter and Magellan flows proceeded smoothly until the launch was scrubbed at T-31 sec.

Source record↗

Trace Contaminant Control for the International Space Station's Node 1- Analysis, Design, and Verification

Trace chemical contaminant generation inside crewed spacecraft cabins is a technical and medical problem that must be continuously evaluated. Although passive control through materials selection and active control by adsorption and catalytic oxidation devices is employed during normal operations of a spacecraft, contaminant buildup can still become a problem. Buildup is particularly troublesome during the stages between the final closure of a spacecraft during ground processing and the time that a crewmember enters for the first time during the mission. Typically, the elapsed time between preflight closure and first entry on orbit for spacecraft such as Spacelab modules was 30 days. During that time, the active contamination control systems are not activated and contaminants can potentially build up to levels which exceed the spacecraft maximum allowable concentrations (SMACs) specified by NASA toxicology experts. To prevent excessively high contamination levels at crew entry, the Spacelab active contamination control system was operated for 53 hours just before launch.

Perry, J. L.↗

Portable computers as companions in space

The paper describes a Space Shuttle Program payload that uses a Macintosh portable computer to gather human performance data for a cursor control device experiment and to test prototype software applications for the Space Station Freedom (SSF). The payload is the first phase of a NASA research and development project to operationally test human-computer interface requirements and crew support applications for an advanced portable computer for the SSF.

Microgravity↗

Tool for simulating space exploration initiative (SEI) operations as a means of obtaining intelligent systems requirements

The Lyndon B. Johnson Space Center Automation and Robotics Division has developed a simulation tool for use in evaluating the feasibility and effectiveness of proposed candidate mission concepts, such as the First Lunar Outpost design reference missions. The tool helps in defining requirements for such missions at the beginning of the definition activities and can be applied iteratively thereafter on a short cycle basis as changes are made to the mission definition. The simulation is applicable to the amount and capabilities of automation and robotics; number and skills of crew; amount of engineering work, science, and in-space materials utilization to be accomplished; impact of provision for maintenance and repair; and flight schedules and manifests. The simulation accounts for supply and demand of resources to accomplish tasks and accounts for elapsed time in carrying out mission processes and tasks.

Erickson, Jon D.↗

Analysis of Apollo 8 Photography and Visual Observations

Apollo 8 was launched from Cape Kennedy, Fla., at 7 :50 a.m., e.s.t., on December 21, 1968. Two hours 50 minutes later, translunar injection was performed; and astronauts Col. Frank Borman, the commander; Capt. James A. Lovell, Jr., the command module pilot; and Maj. (now Lt. Col.) William A. Anders, the lunar module pilot, were on their way to the Moon. The spacecraft was placed in an elliptical lunar orbit at 69 hours 8 minutes after liftoff. After flying two elliptical orbits of 168.5 by 60 nautical miles with an inclination of 12° to the Equator, the spacecraft was placed in a nearly circular orbit of 59.7 by 60.7 nautical miles, in which it remained for eight orbits. At 89 hours 19 minutes, trans earth injection was performed from behind the Moon. A nearly flawless mission was completed on the morning of December 27 when splashdown occurred in the Pacific Ocean after a total elapsed time of 147 hours. Lt. Gen. Sam C. Phillips, the Director of the Apollo Program, announced that such a mission was being considered at a press conference on August 19, 1968. Formal announcement that NASA was preparing Apollo 8 for an orbital flight around the Moon was released to the press on November 12, 1968. The primary purpose of this mission was to further progress toward the goal of landing men on the Moon by gaining operational experience and testing the Apolio systems. However, a great effort was also made to accomplish worthwhile scientific tasks with photography and visual observations by the astronauts. In planning the scientific tasks to be attempted on this mission, it was obvious that one of the prime tasks should be photography of the lunar surface. Such photography would furnish valuable information on the following : 1. Approach topography and landmarks for the early Apollo landings 2. The scientific merit and the roughness of areas for possible follow-on Apollo landings 3. The broad structure and characteristics of the lunar surface During the orbital part of the mission, a major portion of the lunar far side would be in sunlight. Although almost all of the far side of the Moon has been photographed by the automated Lunar Orbiter spacecraft, the photography generally was made with the spacecraft relatively far from the Moon, limiting the Lunar Orbiter photographs to an average resolution of approximately 100 meters. Thus, Apollo photographs of the far side would have much better resolution than existing pictures. Finally, it was recognized that contamination, both as it relates to window fogging (which did occur) and to contamination clouds around t he spacecraft, should be studied for both scientific and operational interests.

Richard J. Allenby↗

Apollo 12 Mission Report

The Apollo 12 mission was the twelfth in a series of flights using Apollo flight hardware and was the second lunar landing. The purpose of the mission was to perform a precise lunar landing and to conduct a specific scientific exploration of a designated landing site in the Ocean of Storms. Since the performance of the entire spacecraft was excellent, this report discusses only the systems performance that significantly differed from that of previous missions. Because they were unique to Apollo 12, the lunar surface experiments, the precision landing operation, and lunar dust contamination are reported in sections 3, 4, and 6, respectively. A complete analysis of all flight data is not possible within the time allowed for preparation of this report. Therefore, report supple­ments will be published for certain Apollo 12 systems analyses, as shown in appendix E. This appendix also lists the current status of all Apollo mission supplements, either published or in preparation. Other supplements will be published as the need is identified. In this report, all actual times prior to earth landing are elapsed time from range zero, established as the integral second before lift-off. Range zero for this mission was 16:22:00 G.m.t., November 14, 1969. Greenwich mean time is used for all times after earth landing as well as for the discussions of the experiments left on the lunar surface. All references to mileage distance are in nautical miles.

Source record↗

Mission description

The accomplishments of the Apollo 17 flight are discussed. The scientific objectives included geological surveying and sampling of materials and surface features in a preselected area of the Taurus-Littrow region, deploying and activating surface experiments, and conducting inflight experiments and photographic tasks during lunar orbit and transearth coast. The individual Apollo 17 experiments and photographic tasks are presented in outline form. Charts are developed to show the major mission events and data collection periods correlated to Greenwich Mean Time and ground elapsed time. Maps of the lunar surface ground track envelope for the Apollo 17 orbiting spacecraft for revolutions one to seventy-five is shown.

Baldwin, R. R.↗