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Results for “TITAN LAUNCH VEHICLE”

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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NASA NTRS1994

Intelsat solar array coupon atomic oxygen flight experiment

A Hughes communications satellite (INTELSAT series) belonging to the INTELSAT Organization was marooned in low-Earth orbit (LEO) on March 14, 1990, following failure of the Titan launch vehicle third stage to separate properly. The satellite, INTELSAT 6, was designed for service in geosynchronous orbit and contains several materials that are potentially susceptible to attack by atomic oxygen. Analysis showed that direct exposure of the silver interconnects in the satellite photovoltaic array to atomic oxygen in LEO was the key materials issue. Available data on atomic oxygen degradation of silver are limited and show high variance, so solar array configurations of the INTELSAT 6 type and individual interconnects were tested in ground-based facilities and during STS-41 (Space Shuttle Discovery, October 1990) as part of the ISAC flight experiment. Several materials for which little or no flight data exist were also tested for atomic oxygen reactivity. Dry lubricants, elastomers, and polymeric and inorganic materials were exposed to an oxygen atom fluence of 1.1 x 10(exp 20) atoms cm(exp 2). Many of the samples were selected to support Space Station Freedom design and decision making. This paper provides an overview of the ISAC flight experiment and a brief summary of results. In addition to new data on materials not before flown, ISAC provided data supporting the decision to rescue INTELSAT 6, which was successfully undertaken in May 1992.

TRT-GENERAL REVIEW↗
NASA NTRS1992

TITAN III/Mars Observer Post-Launch Press Conference

Footage shows George Diller, NASA's Public Affairs Officer, as he introduces the panel members. Speakers includes: William Piotrowski, Program Manager from NASA Headquarters; James Womack, NASA Launch Manager from Kennedy Space Center; John Gibb, TITAN Launch Vehicle Manager from Langley Research Center; Sid Saucier, Transfer Orbit Stage, Director from Marshall Space Flight Center; and David Evans, Mars Observer, Director from the Jet Propulsion Laboratory. The speakers discuss the launch procedures, activities, and some trouble that the Observer is having. The panelists are also seen answering questions from both the audience as well as other NASA Centers.

Source record↗
NASA NTRS1993

Demonstration of the Dynamic Flowgraph Methodology using the Titan 2 Space Launch Vehicle Digital Flight Control System

Dynamic Flowgraph Methodology (DFM) is a new approach developed to integrate the modeling and analysis of the hardware and software components of an embedded system. The objective is to complement the traditional approaches which generally follow the philosophy of separating out the hardware and software portions of the assurance analysis. In this paper, the DFM approach is demonstrated using the Titan 2 Space Launch Vehicle Digital Flight Control System. The hardware and software portions of this embedded system are modeled in an integrated framework. In addition, the time dependent behavior and the switching logic can be captured by this DFM model. In the modeling process, it is found that constructing decision tables for software subroutines is very time consuming. A possible solution is suggested. This approach makes use of a well-known numerical method, the Newton-Raphson method, to solve the equations implemented in the subroutines in reverse. Convergence can be achieved in a few steps.

Yau, M.↗
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 NTRS1990

Final Environmental Impact Statement for the Ulysses Mission (Tier 2)

This Final (Tier 2) Environmental Impact Statement (FEIS) addresses the environmental impacts which may be caused by implementation of the Ulysses mission, a space flight mission to observe the polar regions of the Sun. The proposed action is completion of preparation and operation of the Ulysses spacecraft, including its planned launch at the earliest available launch opportunity on the Space Transportation System (STS) Shuttle in October 1990 or in the backup opportunity in November 1991. The alternative is canceling further work on the mission. The Tier 1 EIS included a delay alternative which considered the Titan 4 launch vehicle as an alternative booster stage for launch in 1991 or later. This alternative was further evaluated and eliminated from consideration when, in November 1988, the U.S. Air Force, which procures the Titan 4, notified that it could not provide a Titan 4 vehicle for the 1991 launch opportunity because of high priority Department of Defense requirements. The Titan 4 launch vehicle is no longer a feasible alternative to the STS/Inertial Upper Stage (IUS)/Payload Assist Module-Special (PAM-S) for the November 1991 launch opportunity. The only expected environment effects of the proposed action are associated with normal launch vehicle operation and are treated elsewhere. The environmental impacts of normal Shuttle launches were addressed in existing NEPA documentation and are briefly summarized. These impacts are limited largely to the near-field at the launch pad, except for temporary stratospheric ozone effects during launch and occasional sonic boom effects near the landing site. These effects were judged insufficient to preclude Shuttle launches. There could also be environmental impacts associated with the accidental release of radiological material during launch, deployment, or interplanetary trajectory injection of the Ulysses spacecraft. Intensive analysis indicates that the probability of release is small. There are no environmental impacts in the no-action alternative; however, the U.S. Government and the European Space Agency would suffer adverse fiscal and programmatic impacts if this alternative were adopted. The scientific benefits of the mission would be delayed and possibly lost. There could be significant impacts on the ability of the U.S. to negotiate international agreements for cooperative space activities.

Source record↗
NASA NTRS1990

Draft Environmental Impact Statement for the Ulysses Mission (Tier 2)

This Draft Environmental Impact Statement (DEIS) addresses the environmental impacts which may be caused by the preparation and operation of the Ulysses spacecraft, including its planned launch on the Space Transportation System (STS) Shuttle and the alternative of canceling further work on the mission. The launch configuration will use the STS/Inertial Upper Stage (IUS)/Payload Assist Module-Special(PAM-S) combination. The Tier 1 EIS included a delay alternative which considered the Titan 4 launch vehicle as an alternative booster stage for launch in 1991 or later. However, the U.S. Air Force, which procures the Titan 4 for NASA, could not provide a Titan 4 vehicle for the 1991 launch opportunity because of high priority Department of Defense requirements. The only expected environmental effects of the proposed action are associated with normal Shuttle launch operations. These impacts are limited largely to the near-field at the launch pad, except for temporary stratospheric ozone effects during launch and occasional sonic boom effects near the landing site. These effects have been judged insufficient to preclude Shuttle launches. In the event of (1) an accident during launch, or (2) reentry of the spacecraft from earth orbit, there are potential adverse health and environmental effects associated with the possible release of plutonium dioxide from the spacecraft's radioisotope thermoelectric generators (RTG). The potential effects considered in this EIS include risks of air and water quality impacts, local land area contamination, adverse health and safety impacts, the disturbance of biotic resources, impacts on wetland areas or areas containing historical sites, and socioeconomic impacts. Intensive analysis of the possible accidents associated with the proposed action are underway and preliminary results indicate small health or environmental risks. The results of a Final Safety Analysis Report will be available for inclusion into the final EIS.

Source record↗
NASA NTRS1988

Draft environmental impact statement for the Galileo Mission (Tier 2)

This Draft Environmental Impact Statement (DEIS) addresses the environmental impacts which may be caused by the preparation and operation of the Galileo spacecraft, including its planned launch on the Space Transportation System (STS) Shuttle and the alternative of canceling further work on the mission. The launch configuration will use the STS/Inertial Upper Stage (IUS)/Payload Assist Module-Special (PAM-S) combination. The Tier 1 EIS included a delay alternative which considered the Titan 4 launch vehicle as an alternative booster stage for launch in 1991 or later. However, the U.S. Air Force, which procures the Titan 4 for NASA, could not provide a Titan 4 vehicle for the 1991 launch opportunity because of high priority Department of Defense requirements. The only expected environmental effects of the proposed action are associated with normal Shuttle launch operations. These impacts are limited largely to the near-field at the launch pad, except for temporary stratospheric ozone effects during launch and occasional sonic boom effects near the landing site. These effects have been judged insufficient to preclude Shuttle launches. In the event of: (1) an accident during launch, or (2) reentry of the spacecraft from earth orbit, there are potential adverse health and environmental effects associated with the possible release of plutonium dioxide from the spacecraft's radioisotope thermoelectric generators (RTG).

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NASA NTRS1976

Baseline meteorological soundings for parametric environmental investigations at Kennedy Space Center and Vandenberg Air Force Base

Meteorological soundings representative of the atmospheric environment at the Kennedy Space Center, Florida and Vandenberg Air Force Base, California, are presented. Synthetic meteorological soundings at Kennedy Space Center, including fall, spring, and a sea breeze, and at Vandenberg Air Force Base (sea breeze with low and high level inversion and stationary upper level troughs) are shown. Soundings of frontal passages are listed. The Titan launch soundings at Kennedy Space Center present a wide range of meteorological conditions, both seasonal and time of day variations. The meteorological data input of altitude, wind speed, wind direction, temperature, and pressure may be used as meteorological inputs for the NASA/MSFC Multilayer Diffusion Model or other models to obtain quantitative estimates of effluent concentrations associated with the potential emission of major combustion products in the lower atmosphere to simulate actual launches of space vehicles. The Titan launch soundings are also of value in terms of rocket effluent measurements for analysis purposes.

Susko, M.↗
NASA NTRS1979

Launch vehicle effluent measurements during the May 12, 1977, Titan 3 launch at Air Force Eastern Test Range

Airborne effluent measurements and cloud physical behavior for the May 21, 1977, Titan 3 launch from the Air Force Eastern Test Range, Fla. are presented. The monitoring program included airborne effluent measurements in situ in the launch cloud, visible and infrared photography of cloud growth and physical behavior, and limited surface collection of rain samples. Airborne effluent measurements included concentrations of HCl, NO, NOx, and aerosols as a function of time in the exhaust cloud. For the first time in situ particulate mass concentration and aerosol number density were measured as a function of time and size in the size range of 0.05 to 25 micro meters diameter. Measurement results were similar to those of earlier launch monitorings. Maximum HCl and NOx concentrations ranged from 10 ppm and 500 ppb, respectively, several minutes after launch to about 1 ppm and 100 ppb at 45 minutes after launch.

Gregory, G. L.↗
NASA NTRS1979

Launch vehicle effluent measurements during the August 20, 1977, Titan 3 launch at Air Force Eastern Test Range

Airborne effluent measurements within the launch cloud and visible and infrared measurements of cloud physical behavior are discussed. Airborne effluent measurements include concentrations of HCl, Cl2, NO, NOX, and particulates as a function of time during each sampling pass through the exhaust cloud. The particle size distribution was measured for each pass through the cloud. Mass concentration as a function of particle diameter was measured over the size range of 0.05- to 25 micron diameter, and particle number density was measured as a function of diameter over a size range of 0.5 to 7.5 micron. Effluent concentrations in the cloud ranged from about 30 ppm several minutes after launch to about 1 to 2 ppm at 100 minutes. Maximum Cl2 concentrations were about 40 to 55 ppb and by 20 minutes were less than 1.0 ppb. A tabulated listing of the airborne data is given in the appendix. Usable cloud imaging data were limited to the first 16 minutes after launch.

Woods, D. C.↗
NASA NTRS1979

Launch vehicle effluent measurements during the September 5, 1977, Titan 3 launch at Air Force eastern test range

Airborne effluent measurements and cloud physical behavior data are presented. The monitoring program included airborne effluent measurements in situ in the launch cloud, visible and infrared photography of cloud growth and physical behavior, and limited surface collection of rain samples. Effluent measurements included concentrations of HCl, Cl2, NO, nitric oxide, and particles as a function of time in the exhaust cloud. In situ particle mass concentration and number density were measured as a function of time and size in the range of 0.05 micron m to 30 micron m diameter. Measurement results were similar to those of previous launch monitorings. Maximum HCl and nitric oxide concentrations of Cl2 were maximum about 2 minutes after launch and by 10 to 15 minutes had decayed to less than 10 ppb (detection limit). Particle measurements showed most of the particles present to be below about 3-micron m diameter. Postlaunch analyses of collected particle samples showed significant amounts of Al (some cases Cl) from about 3-micron m to 0.04-micron m diameter.

Wornom, D. E.↗
NASA NTRS1973

Ballistic mode Mercury orbiter missions.

The MVM'73 Mercury flyby mission will initiate exploration of this unique planet. No firm plans for follow-on investigations have materialized due to the difficult performance requirements of the next logical step, an orbiter mission. Previous investigations of ballistic mode flight opportunities have indicated requirements for a Saturn V class launch vehicle. Consequently, most recent effort has been oriented to use of solar electric propulsion. More comprehensive study of the ballistic flight mode utilizing Venus gravity-assist has resulted in identification of timely high-performance mission opportunities compatible with programmed launch vehicles and conventional spacecraft propulsion technologies. A likely candidate for an initial orbiter mission is a 1980 opportunity which offers net orbiter spacecraft mass of about 435 kg with the Titan IIIE/Centaur launch vehicle and single stage solid propulsion for orbit insertion.

Hollenbeck, G. R.↗
NASA NTRS1975

Mars sample return by orbital rendezvous

The return of Mars surface samples by the orbital rendezvous mode would require smaller, lighter vehicles than the direct return mode and would limit the exposure of the earth return systems to possible contamination by Martian biota. The reference mission considered would use a single Titan IIIE/Centaur launch vehicle, carry a 1 kg sample, and be launched in late 1981. The spacecraft would consist of an orbiter, a Mars lander, a Mars ascent vehicle, an earth return vehicle, and an earth entry capsule. The Mars ascent vehicle is the only all-new vehicle required, the others being derived from Viking and other programs. A key consideration is the feasibility of automatic rendezvous and sample transfer between the ascent vehicle and orbiter in Mars orbit. The maneuvers required and the maneuver algorithm developed in a NASA feasibility study are described in detail. The possibility of international cooperation on a Mars orbital rendezvous mission is discussed.

Moore, J. W.↗
NASA NTRS1993

Meeting with a Majestic Giant: The Cassini Mission to Saturn

In October of 1997, a two-story-tall robotic spacecraft will begin a journey of many years to the vast and exiting realm of Saturn. With a mass of roughly 2,500 kilograms (5,510 pounds) of dry hardware and 3,000 kilograms (6,615 pounds) of propellant, it needs a boost from the Titan IV/Centaur launch vehicle and several planetary gravity assists. Both are needed if Cassini is to reach Saturn with sufficient propellant to brake into Saturn orbit and accomplish its mission: to deliver a European-built probe to the large, hazy moon Titan, and then tour the saturnian system for nearly four years. The Cassini mission has been undertaken by NASA, the European Space Agency (ESA) and the Italian Space Agency (ASI). It is named in honor of the French-Italian astronomer Jean Dominique Cassini, who discovered the prominent gap in Saturn's main rings (now called the Cassini division), as well as the icy moons Iapetus, Rhea, Tethys and Dione...

Kohlhase, C.↗
NASA NTRS1993

Ascent abort capability for the HL-20

The HL-20 has been designed with the capability for rescue of the crew during all phases of powered ascent from on the launch pad until orbital injection. A launch-escape system, consisting of solid rocket motors located on the adapter between the HL-20 and the launch vehicle, provides the thrust that propels the HL-20 to a safe distance from a malfunctioning launch vehicle. After these launch-escape motors have burned out, the adapter is jettisoned and the HL-20 executes one of four abort modes. In three abort modes - return-to-launch-site, transatlantic-abort-landing, and abort-to-orbit - not only is the crew rescued, but the HL-20 is recovered intact. In the ocean-landing-by-parachute abort mode, which occurs in between the return-to-launch-site and the transatlantic-abort-landing modes, the crew is rescued, but the HL-20 would likely sustain damage from the ocean landing. This paper describes the launch-escape system and the four abort modes for an ascent on a Titan III launch vehicle.

Naftel, J. C.↗