Results for “LAUNCH DATE” — Page 4

NASA NTRS2015

Multi-Objective Hybrid Optimal Control for Multiple-Flyby Interplanetary Mission Design Using Chemical Propulsion

The customer (scientist or project manager) most often does not want just one point solution to the mission design problem Instead, an exploration of a multi-objective trade space is required. For a typical main-belt asteroid mission the customer might wish to see the trade-space of: Launch date vs. Flight time vs. Deliverable mass, while varying the destination asteroid, planetary flybys, launch year, etcetera. To address this question we use a multi-objective discrete outer-loop which defines many single objective real-valued inner-loop problems.

Design↗

NASA NTRS1963

Satellite Situation Report

Satellite situation report - name, source, launch date, nodal period, inclination, apogee, perigee, and transmitting frequency

SATELLITE LAUNCHING↗

NASA NTRS1983

Interplanetary trajectory design for the Galileo mission

The Galileo mission has been reprogrammed to use a direct earth-Jupiter trajectory with a May 1986 launch date and with arrival at Jupiter occurring in mid-1988. Within the constraints of Shuttle/Centaur launch vehicle capability and total spacecraft mass and performance, optimal broken-plane trajectories are generated, and the region of positive propellant margin in the launch/arrival space is determined. Mission constraints are used to define a launch/arrival strategy. It is also shown that a close flyby of any one of several asteroids or a comet is possible on the interplanetary transfer with minimal impact on mission performance.

Damario, L. A.↗

NASA NTRS2004

A New Spin

The trouble was that the shuttle was still under development when that schedule was set. As time went on, the Shuttle had problems with its high pressure turbines, thermal protection tiles, engines, and more. The early launch dates had to be scrapped. NASA Headquarters told us, "We re going to delay your launch two years to allow more time for the Shuttle development to take place. You can slow your development accordingly." Right off the bat, we looked into the celestial mechanics and how they would affect us. The difficulty in launching a spacecraft to Jupiter changes on a year-to- year basis, in a cyclical pattern that repeats about every ten or twelve years. In order to achieve the velocity needed to get from low earth orbit to Jupiter, an upper stage is required in the Shuttle. For the 1982 launch the upper stage was adequate, but it could not provide the velocity we would need in 1984. This meant we would have to separate the Galileo probe from the Galileo orbiter before launch and put each of them on separate Shuttles with separate upper stages. When we told the folks at Headquarters this, they told us, "Okay we'll give you two Shuttle launches."

Casani, John↗

NASA NTRS1983

Missions to the asteroid Anteros and the space of true anomalies

Three opportunities for missions to rendezvous ballistically with the earth-crossing asteroid Anteros are studied to illustrate the requirements for a trip to a near-earth minor planet. The rationale, sample payload, spacecraft requirements and trajectory characteristics of these opportunities are typical of a rendezvous mission to an accessible near-earth object. Round trip ballistic trajectories to return small samples of the asteroid with launch dates between 1985 and 2000 are also presented. Contours of minimum total delta drawn in the space of launch and arrival true anomalies, given the designation Prime Rib curves, are introduced as a useful tool for mission design.

Hulkower, N. D.↗

NASA NTRS1992

Delta II Geotail -- 1st Stage and Solid Motor Booster Erection

The Geotail mission's goal was to investigate the structure and dynamics of the geomagnetic tail that extends on the nightside of the Earth. The launch date was July 24, 1992. This video shows the Delta II on the pad, being prepared for the launch. The first stage and the solid motor booster are shown being moved into place on the rocket.

Source record↗

NASA NTRS2016

The Exploration of Mars Launch and Assembly Simulation

Advancing human exploration of space beyond Low Earth Orbit, and ultimately to Mars, is of great interest to NASA, other organizations, and space exploration advocates. Various strategies for getting to Mars have been proposed. These include NASA's Design Reference Architecture 5.0, a near-term flyby of Mars advocated by the group Inspiration Mars, and potential options developed for NASA's Evolvable Mars Campaign. Regardless of which approach is used to get to Mars, they all share a need to visualize and analyze their proposed campaign and evaluate the feasibility of the launch and on-orbit assembly segment of the campaign. The launch and assembly segment starts with flight hardware manufacturing and ends with final departure of a Mars Transfer Vehicle (MTV), or set of MTVs, from an assembly orbit near Earth. This paper describes a discrete event simulation based strategic visualization and analysis tool that can be used to evaluate the launch campaign reliability of any proposed strategy for exploration beyond low Earth orbit. The input to the simulation can be any manifest of multiple launches and their associated transit operations between Earth and the exploration destinations, including Earth orbit, lunar orbit, asteroids, moons of Mars, and ultimately Mars. The simulation output includes expected launch dates and ascent outcomes i.e., success or failure. Running 1,000 replications of the simulation provides the capability to perform launch campaign reliability analysis to determine the probability that all launches occur in a timely manner to support departure opportunities and to deliver their payloads to the intended orbit. This allows for quantitative comparisons between alternative scenarios, as well as the capability to analyze options for improving launch campaign reliability. Results are presented for representative strategies.

Cates, Grant↗

NASA NTRS2016

The Exploration of Mars Launch and Assembly Simulation

Advancing human exploration of space beyond Low Earth Orbit, and ultimately to Mars, is of great interest to NASA, other organizations, and space exploration advocates. Various strategies for getting to Mars have been proposed. These include NASA's Design Reference Architecture 5.0, a near-term flyby of Mars advocated by the group Inspiration Mars, and potential options developed for NASA's Evolvable Mars Campaign. Regardless of which approach is used to get to Mars, they all share a need to visualize and analyze their proposed campaign and evaluate the feasibility of the launch and on-orbit assembly segment of the campaign. The launch and assembly segment starts with flight hardware manufacturing and ends with final departure of a Mars Transfer Vehicle (MTV), or set of MTVs, from an assembly orbit near Earth. This paper describes a discrete event simulation based strategic visualization and analysis tool that can be used to evaluate the launch campaign reliability of any proposed strategy for exploration beyond low Earth orbit. The input to the simulation can be any manifest of multiple launches and their associated transit operations between Earth and the exploration destinations, including Earth orbit, lunar orbit, asteroids, moons of Mars, and ultimately Mars. The simulation output includes expected launch dates and ascent outcomes i.e., success or failure. Running 1,000 replications of the simulation provides the capability to perform launch campaign reliability analysis to determine the probability that all launches occur in a timely manner to support departure opportunities and to deliver their payloads to the intended orbit. This allows for quantitative comparisons between alternative scenarios, as well as the capability to analyze options for improving launch campaign reliability. Results are presented for representative strategies.

Cates, Grant↗

NASA NTRS2007

Validation of the Pulmonary Function System for Use on the International Space Station

Aerobic deconditioning occurs during long duration space flight despite the use of exercise countermeasures (Convertino, 1996). As a part of International Space Station (ISS) medical operations, periodic tests designed to estimate aerobic capacity are performed to track changes in aerobic fitness and to determine the effectiveness of exercise countermeasures. These tests are performed prior to, during, and after missions of greater than 30 days in duration. Crewmembers selected for missions aboard the ISS perform a graded exercise test on a cycle ergometer approximately 270 days prior to their scheduled launch date in order to measure peak oxygen consumption (VO2PK) and peak heart rate (HRpk). Approximately 30 to 45 days prior to launch, crewmembers perform a submaximal cycle ergometer test at work rates set to elicit 25, 50 and 75% of their pre-flight VO2PK. This test, known as the Periodic Fitness Evaluation (PFE), serves as a baseline measure to which subsequent in-and post-flight exercise tests are compared. While onboard the ISS, crewmembers are normally scheduled to perform the PFE beginning with flight day (FD) 14 and every 30 days thereafter. The PFE is also conducted 5 and 30 days following flight. Using PFE data, aerobic fitness is estimated by quantifying the VO2 vs. HR relationship using linear regression and calculating the VO2 that would occur at the crewmember s previously measured HRpk. Currently, for data collected during flight, this technique assumes that the pre- vs. in-flight oxygen consumption per given cycle workload is similar. However, the validity of this assumption is based upon a sparse amount of data collected during the Skylab era (Michel, et al. 1977). The method of using heart rate and cycle ergometer work rates has been used to estimate aerobic fitness in normal gravity (Astrand and Ryhming, 1954; Lee, 1993). Due to spaceflight induced physiological alterations, such as shifts in extracellular fluid (e.g. plasma) volume, this method may not be valid during space flight. In addition, the ergometer onboard ISS is vibration-isolated and moves with the astronaut s application of force into the pedals. The effect of this movement on the VO2 of cycle exercise on ISS has not been quantified.

McCleary, Frank A.↗

NASA NTRS2011

Power and Propulsion System Design for Near-Earth Object Robotic Exploration

Near-Earth Objects (NEOs) are exciting targets for exploration; they are relatively easy to reach but relatively little is known about them. With solar electric propulsion, a vast number of interesting NEOs can be reached within a few years and with extensive flexibility in launch date. An additional advantage of electric propulsion for these missions is that a spacecraft can be small, enabling a fleet of explorers launched on a single vehicle or as secondary payloads. Commercial, flight-proven Hall thruster systems have great appeal based on their performance and low cost risk, but one issue with these systems is that the power processing units (PPUs) are designed for regulated spacecraft power architectures which are not attractive for small NEO missions. In this study we consider the integrated design of power and propulsion systems that utilize the capabilities of existing PPUs in an unregulated power architecture. Models for solar array and engine performance are combined with low-thrust trajectory analyses to bound spacecraft design parameters for a large class of NEO missions, then detailed array performance models are used to examine the array output voltage and current over a bounded mission set. Operational relationships between the power and electric propulsion systems are discussed, and it is shown that both the SPT-100 and BPT-4000 PPUs can perform missions over a solar range of 0.7 AU to 1.5 AU - encompassing NEOs, Venus, and Mars - within their operable input voltage ranges. A number of design trades to control the array voltage are available, including cell string layout, array offpointing during mission operations, and power draw by the Hall thruster system.

Near Earth Object (NEO)↗

NASA NTRS2012

Progress Towards a 2012 Landsat Launch

The Landsat Data Continuity Mission (LDCM) is on schedule for a December 2012 launch date. The mission is being managed by an interagency partnership between NASA and the U.S. Geological Survey (USGS). NASA leads the development and launch of the satellite observatory while leads ground system development. USGS will assume responsibility for operating the satellite and for collecting, archiving, and distributing the LDCM data following launch. When launched the satellite will carry two sensors into orbit. The Operational Land Imager (OLI) will collect data for nine shortwave spectral bands with a spatial resolution of 30 m (with a 15 m panchromatic band). The Thermal Infrared Sensor (TIRS) will coincidently collect data for two thermal infrared bands with a spatial resolution of 100 m. The OLI is fully assembled and tested and has been shipped by it?s manufacturer, Ball Aerospace and Technology Corporation, to the Orbital Sciences Corporation (Orbital) facility where it is being integrated onto the LDCM spacecraft. Pre-launch testing indicates that OLI will meet all performance specification with margin. TIRS is in development at the NASA Goddard Space Flight Center (GSFC) and is in final testing before shipping to the Orbital facility in January, 2012. The ground data processing system is in development at the USGS Earth Resources Observation and Science (EROS) Center. The presentation will describe the LDCM satellite system, provide the status of system development, and present prelaunch performance data for OLI and TIRS. The USGS has committed to renaming the satellite as Landsat 8 following launch.

Irons, Jim↗

NASA NTRS1974

Space Trajectory Error Analysis Program (STEAP) for halo orbit missions. Volume 1: Analytic and user's manual

Development, test, conversion, and documentation of computer software for the mission analysis of missions to halo orbits about libration points in the earth-sun system is reported. The software consisting of two programs called NOMNAL and ERRAN is part of the Space Trajectories Error Analysis Programs (STEAP). The program NOMNAL targets a transfer trajectory from Earth on a given launch date to a specified halo orbit on a required arrival date. Either impulsive or finite thrust insertion maneuvers into halo orbit are permitted by the program. The transfer trajectory is consistent with a realistic launch profile input by the user. The second program ERRAN conducts error analyses of the targeted transfer trajectory. Measurements including range, doppler, star-planet angles, and apparent planet diameter are processed in a Kalman-Schmidt filter to determine the trajectory knowledge uncertainty. Execution errors at injection, midcourse correction and orbit insertion maneuvers are analyzed along with the navigation uncertainty to determine trajectory control uncertainties and fuel-sizing requirements. The program is also capable of generalized covariance analyses.

Byrnes, D. V.↗

NASA NTRS1974

Space Trajectory Error Analysis Program (STEAP) for halo orbit missions. Volume 2: Programmer's manual

The six month effort was responsible for the development, test, conversion, and documentation of computer software for the mission analysis of missions to halo orbits about libration points in the earth-sun system. The software consisting of two programs called NOMNAL and ERRAN is part of the Space Trajectories Error Analysis Programs. The program NOMNAL targets a transfer trajectory from earth on a given launch date to a specified halo orbit on a required arrival date. Either impulsive or finite thrust insertion maneuvers into halo orbit are permitted by the program. The transfer trajectory is consistent with a realistic launch profile input by the user. The second program ERRAN conducts error analyses of the targeted transfer trajectory. Measurements including range, doppler, star-planet angles, and apparent planet diameter are processed in a Kalman-Schmidt filter to determine the trajectory knowledge uncertainty.

Byrnes, D. V.↗

NASA NTRS2004

[Activities of System Studies and Simulation, Inc.]

Contents include the following: 1. Launch Vehicle Interface Work Performed: a. S3 provided to KSC the new launch inclination targets needed for the April '04 launch date. 2. Prelaunch operations work performed: a. S3 updated the staffing plan for MSFC on-console personnel\during the Final Countdown prior to launch. 3. Software Assessment Work Performed: a. S3 evaluated and recommended approval for Program Control Board (PCB) proposed change 649 for ground software changes, as well as change 650 and 650A for Stored Program Commands. 4. Education and Public Outreach Work Performed: a. S3 continues to coordinate the effort for the design and fabrication of scale models of the GP-3 for use at the launch site, education forums, and management/technical briefings. S3 also prepared a Change Request for additional funds needed for fabrication of additional scale models. S3 drafted the planned uses of these models, including the possibility of participation in the Boston, MA showings of the traveling Einstein Exhibit. 5. Program Management Support Work Performed: a. S3 prepared the input for and closed three MSFC Centerwide Action Item Tracking Systems (CAITS) actions during this period.

Source record↗

NASA NTRS2012

Systems Engineering and Technology Considerations of a Mars Ascent Vehicle

A Mars Ascent Vehicle (MAV) systems engineering study is underway to define the driving requirements, system architecture, major risks, and required technology developments to support the launch of a rock core sample to a specified delivery orbit for later retrieval and return to Earth. The proposed MAV would essentially be a small-scale launch vehicle, the first of its kind to be launched autonomously from another planet. The MAV would be a flight element of the proposed Mars Sample Return (MSR) campaign architecture, which currently assumes a 2018 launch of the sample caching mission and a 2024 (Earth) launch date of the MAV and lander, with arrival on Mars in 2025. After 9 months on the surface the MAV would be erected and launched to a specified delivery orbit. In the delivery orbit it would release its payload, a 5 kg sphere containing the rock core sample. An orbiter would rendezvous and capture the payload, returning it to Earth a year later.

entry, descent, and landing (EDL)↗

NASA NTRS2011

Ares I-X Range Safety Analyses Overview

Ares I-X was the first test flight of NASA's Constellation Program's Ares I Crew Launch Vehicle designed to provide manned access to low Earth orbit. As a one-time test flight, the Air Force's 45th Space Wing required a series of Range Safety analysis data products to be developed for the specified launch date and mission trajectory prior to granting flight approval on the Eastern Range. The range safety data package is required to ensure that the public, launch area, and launch complex personnel and resources are provided with an acceptable level of safety and that all aspects of prelaunch and launch operations adhere to applicable public laws. The analysis data products, defined in the Air Force Space Command Manual 91-710, Volume 2, consisted of a nominal trajectory, three sigma trajectory envelopes, stage impact footprints, acoustic intensity contours, trajectory turn angles resulting from potential vehicle malfunctions (including flight software failures), characterization of potential debris, and debris impact footprints. These data products were developed under the auspices of the Constellation's Program Launch Constellation Range Safety Panel and its Range Safety Trajectory Working Group with the intent of beginning the framework for the operational vehicle data products and providing programmatic review and oversight. A multi-center NASA team in conjunction with the 45th Space Wing, collaborated within the Trajectory Working Group forum to define the data product development processes, performed the analyses necessary to generate the data products, and performed independent verification and validation of the data products. This paper outlines the Range Safety data requirements and provides an overview of the processes established to develop both the data products and the individual analyses used to develop the data products, and it summarizes the results of the analyses required for the Ares I-X launch.

Starr, Brett R.↗

NASA NTRS2011

Ares I-X Range Safety Flight Envelope Analysis

Ares I-X was the first test flight of NASA's Constellation Program's Ares I Crew Launch Vehicle designed to provide manned access to low Earth orbit. As a one-time test flight, the Air Force's 45th Space Wing required a series of Range Safety analysis data products to be developed for the specified launch date and mission trajectory prior to granting flight approval on the Eastern Range. The range safety data package is required to ensure that the public, launch area, and launch complex personnel and resources are provided with an acceptable level of safety and that all aspects of prelaunch and launch operations adhere to applicable public laws. The analysis data products, defined in the Air Force Space Command Manual 91-710, Volume 2, consisted of a nominal trajectory, three sigma trajectory envelopes, stage impact footprints, acoustic intensity contours, trajectory turn angles resulting from potential vehicle malfunctions (including flight software failures), characterization of potential debris, and debris impact footprints. These data products were developed under the auspices of the Constellation's Program Launch Constellation Range Safety Panel and its Range Safety Trajectory Working Group with the intent of beginning the framework for the operational vehicle data products and providing programmatic review and oversight. A multi-center NASA team in conjunction with the 45th Space Wing, collaborated within the Trajectory Working Group forum to define the data product development processes, performed the analyses necessary to generate the data products, and performed independent verification and validation of the data products. This paper outlines the Range Safety data requirements and provides an overview of the processes established to develop both the data products and the individual analyses used to develop the data products, and it summarizes the results of the analyses required for the Ares I-X launch.

Starr, Brett R.↗

NASA NTRS1979

STS utilization planning

The paper shows how NASA plans to utilize the Shuttle to carry a wide variety of payloads. Topics covered include: Shuttle Transportation System (STS) description and capabilities, payload integration and manifesting payload business today, and NASA-JSC user services. Consideration is given to the flight assignment process and flight manifest which is a listing of flights with their attendant launch dates, payload names, days on-orbit, altitude, inclination, and crew required. Additional data such as weight, center of gravity, payload-on-orbit capability, and launch windows are checked to see if the manifest is still valid. The STS traffic model is described along with the small self-contained payloads.

Bauch, G. T.↗