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The analysis of the gross aspects of the flight mechanics for lunar trajectories from orbital launch is presented. An attempt is made to define the launch requirements from an earth orbit, the geometrical relationships existing between an Orbital Launch Facility in earth orbit and the moon in its orbit, the departure trajectory sensitivity to energy and time requirements, the launch-on-time problem and the orbital "escape window".
The Sounding of the Atmosphere using Broadband Emission Radiometry (SABER) experiment was launched onboard the TIMED satellite in December, 2001. SABER is designed to provide measurements of the key radiative and chemical sources and sinks of energy in the mesosphere and lower thermosphere (MLT). SABER measures Earth limb emission in 10 broadband radiometer channels ranging from 1.27 micrometers to 17 micrometers. Measurements are made both day and night over the latitude range from 54 deg. S to 87 deg. N with alternating hemisphere coverage every 60 days. In this paper we concentrate on retrieved profiles of kinetic temperature (T(sub k)) and CO2 volume mixing ratio (vmr), inferred from SABER-observed 15 micrometer and 4.3 micrometer limb emissions, respectively. SABER-measured limb radiances are in non-local thermodynamic equilibrium (non-LTE) in the MLT region. The complexity of non-LTE radiation transfer combined with the large volume of data measured by SABER requires new retrieval approaches and radiative transfer techniques to accurately and efficiently retrieve the data products. In this paper we present the salient features of the coupled non-LTE T(sub k)/CO2 retrieval algorithm, along with preliminary results.
In the last decade, improvements in computer technology have provided new 'tools' for controlling and monitoring critical missile systems. In this connection, computers have gradually taken a large role in monitoring all flights and ground systems on the Atlas Centaur. The wide body Centaur which will be launched in the Space Shuttle Cargo Bay will use computers to an even greater extent. It is planned to use the wide body Centaur to boost the Galileo spacecraft toward Jupiter in 1985. The critical systems which must be monitored prior to liftoff are examined. Computers have now been programmed to monitor all critical parameters continuously. At this time, there are two separate computer systems used to monitor these parameters.
The current design of the Space Shuttle Main Engine has passed 11 certification cycles, amassed approximately a quarter million seconds of engine test time in 1200 tests and successfully launched the Space Shuttle 17 times of 51 engine launches through May 1985. Building on this extensive background, two development programs are underway at Rocketdyne to improve the flow of hot gas through the powerhead and evaluate the changes to increase the performance margins in the engine. These two programs, called Phase II+ and Technology Test Bed Precursor program are described. Phase II+ develops a two-tube hot-gas manifold that improves the component environment. The Precursor program will evaluate a larger throat main combustion chamber, conduct combustion stability testing of a baffleless main injector, fabricate an experimental weld-free heat exchanger tube, fabricate and test a high pressure oxidizer turbopump with an improved inlet, and develop and test methods for reducing temperature transients at start and shutdown.
The launch vehicle stages for SL-2, the CSM experiments, and their associated support equipment are reported. The performance of KSC systems in support of processing and launch of the SL-2 are described along with major processing events for each launch vehicle stage, the spacecraft, and the general experiments of the SL-2 S/V. The final countdown and hold times are noted and a summary of the launch vehicle is included. The weather conditions at launch time and the range support activities are given.
NASA's Space Launch Initiative (SLI) is an effort to research and develop the technologies needed to build a second-generation reusable launch vehicle. It is required that this new launch vehicle be 100 times safer and 10 times cheaper to operate than current launch vehicles. Part of the SLI includes the development of reusable composite and metallic cryotanks. The size of these reusable tanks is far greater than anything ever developed and exceeds the design limits of current manufacturing tools. Several design and manufacturing approaches have been formulated, but many factors must be weighed during the selection process. Among these factors are tooling reachability, cycle times, feasibility, and facility impacts. The manufacturing process simulation capabilities available at NASA's Marshall Space Flight Center have played a key role in down selecting between the various manufacturing approaches. By creating 3-D manufacturing process simulations, the varying approaches can be analyzed in a virtual world before any hardware or infrastructure is built. This analysis can detect and eliminate costly flaws in the various manufacturing approaches. The simulations check for collisions between devices, verify that design limits on joints are not exceeded, and provide cycle times which aid in the development of an optimized process flow. In addition, new ideas and concerns are often raised after seeing the visual representation of a manufacturing process flow. The output of the manufacturing process simulations allows for cost and safety comparisons to be performed between the various manufacturing approaches. This output helps determine which manufacturing process options reach the safety and cost goals of the SLI.
NASA's Space Launch Initiative (SLI) is an effort to research and develop the technologies needed to build a second-generation reusable launch vehicle. It is required that this new launch vehicle be 100 times safer and 10 times cheaper to operate than current launch vehicles. Part of the SLI includes the development of reusable composite and metallic cryotanks. The size of these reusable tanks is far greater than anything ever developed and exceeds the design limits of current manufacturing tools. Several design and manufacturing approaches have been formulated, but many factors must be weighed during the selection process. Among these factors are tooling reachability, cycle times, feasibility, and facility impacts. The manufacturing process simulation capabilities available at NASA.s Marshall Space Flight Center have played a key role in down selecting between the various manufacturing approaches. By creating 3-D manufacturing process simulations, the varying approaches can be analyzed in a virtual world before any hardware or infrastructure is built. This analysis can detect and eliminate costly flaws in the various manufacturing approaches. The simulations check for collisions between devices, verify that design limits on joints are not exceeded, and provide cycle times which aide in the development of an optimized process flow. In addition, new ideas and concerns are often raised after seeing the visual representation of a manufacturing process flow. The output of the manufacturing process simulations allows for cost and safety comparisons to be performed between the various manufacturing approaches. This output helps determine which manufacturing process options reach the safety and cost goals of the SLI. As part of the SLI, The Boeing Company was awarded a basic period contract to research and propose options for both a metallic and a composite cryotank. Boeing then entered into a task agreement with the Marshall Space Flight Center to provide manufacturing simulation support. This paper highlights the accomplishments of this task agreement, while also introducing the capabilities of simulation software.
Propellant will be more than 85% of the mass that needs to be lofted into Low Earth Orbit (LEO) in the planned program of Exploration of the Moon, Mars, and beyond. This paper describes a possible means for launching thousands of tons of propellant per year into LEO at a cost 15 to 30 times less than the current launch cost per kilogram. The basic idea is to mass-produce very simple, small and relatively low-performance rockets at a cost per kilogram comparable to automobiles, instead of the ~25X greater cost that is customary for current launch vehicles that are produced in small quantities and which are manufactured with performance near the limits of what is possible. These small, simple rockets can reach orbit because they are launched above ~95% of the atmosphere, where the drag losses even on a small rocket are acceptable, and because they can be launched nearly horizontally with very simple guidance based primarily on spin-stabilization. Launching above most of the atmosphere is accomplished by winching the rocket up a tether to a balloon. A fuel depot in equatorial orbit passes over the launch site on every orbit (approximately every 90 minutes). One or more rockets can be launched each time the fuel depot passes overhead, so the launch rate can be any multiple of 6000 small rockets per year, a number that is sufficient to reap the benefits of mass production.
Time-optimal control theory applied to launch vehicle regulation and the newton-raphson optimization method
In the rocket triggered lightning experiment of fiscal 1989, researchers observed electromagnetic field changes and UHF electromagnetic radiation accompanying rocket triggered lightning. It was found that no rapid changes corresponding to the return stroke of natural lightning were observed in the electric field changes accompanying rocket triggered lightning. However, continuous currents were present. In the case of rocket triggered lightning to the tower, electromagnetic field changes corresponding to the initiation of triggered lightning showed a bipolar pulse of a relatively large amplitude. In contrast, the rocket triggered lightning to the ground did not have such a bipolar pulse. The UHF radiation accompanying the rocket triggered lightning preceded the waveform portions corresponding to the first changes in electromagnetic fields. The number of isolated pulses in the UHF radiation showed a correlation with the time duration from rocket launching up to triggered lightning. The time interval between consecutive isolated pulses tended to get shorter with the passage of time, just like the stepped leaders of natural lightning.
Bone biopsies indicate that during immobilization bone marrow adipose tissue increases while the functional cellular fraction decreases. One objective of our Spacelab flight experiment was to determine, using in vivo volume-localized magnetic resonance spectroscopy (VLMRS), whether bone marrow composition was altered by space flight. Four crew members of a 17 day Spacelab mission participated in the experiment. The apparent cellular fraction and transverse relaxation time (T2) were determined twice before launch and at several times after flight. Immediately after flight, no significant change in the cellular fraction was found. However, the T2 of the cellular, but not the fat component increased following flight, although to a variable extent, in all crew members with a time course for return to baseline lasting several months. The T2 of seven control subjects showed no significant change. Although these observations may have several explanations, it is speculated that the observed T2 changes might reflect increased marrow osteoblastic activity during recovery from space flight.
Presentation Agenda: Typical Payload Development Process, Problem, Solution Requirements, Methodology, Time Domain Solution, Launch Vehicle FEM, Heavy Payload FEM, Problem Size Comparison, Frequency Domain Validation, Time Domain Validation
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PatCon code was developed to help mission designers run trade studies on launch and arrival times for any given planet. Initially developed in Fortran, the required inputs included launch date, arrival date, and other orbital parameters of the launch planet and arrival planets at the given dates. These parameters include the position of the planets, the eccentricity, semi-major axes, argument of periapsis, ascending node, and inclination of the planets. With these inputs, a patched conic approximation is used to determine the trajectory. The patched conic approximation divides the planetary mission into three parts: (1) the departure phase, in which the two relevant bodies are Earth and the spacecraft, and where the trajectory is a departure hyperbola with Earth at the focus; (2) the cruise phase, in which the two bodies are the Sun and the spacecraft, and where the trajectory is a transfer ellipse with the Sun at the focus; and (3) the arrival phase, in which the two bodies are the target planet and the spacecraft, where the trajectory is an arrival hyperbola with the planet as the focus.
The Space Launch System (SLS) has a co-manifested payload capability that will grow over time as the launch vehicle matures and planned upgrades are implemented. The final configuration is planned to be capable of inserting a payload greater than 10 metric tons (mt) into a trans-lunar injection trajectory along with the crew in the Orion capsule and its service module. The co-manifested payload is located below the Orion and its service module in a 10 m high fairing similar to the way the Saturn launch vehicle carried the lunar lander below the Apollo command and service modules. Various approaches that utilize this comanifested payload capability to build up infrastructure in deep space have been explored in support of future asteroid, lunar, and Mars mission scenarios. This paper reports on the findings of the Advanced Concepts Office study team at NASA Marshall Space Flight Center (MSFC) working with the Advanced Exploration Systems Program on the Exploration Augmentation Module Project. It includes some of the possible options for habitation in the co-manifested payload volume of the SLS. Findings include a set of module designs that can be developed in 10 mt increments to support these co-manifested payload missions along with a comparison of this approach to a large-module payload flight configuration for the SLS.
The SAM-2 will fly aboard the Nimbus-G satellite for launch in the fall of 1978 and measure stratospheric vertical profiles of aerosol extinction in high latitude bands. The plan gives details of the location and times for the simultaneous satellite/correlative measurements for the nominal launch time, the rationale and choice of the correlative sensors, their characteristics and expected accuracies, and the conversion of their data to extinction profiles. The SAM-2 expected instrument performance and data inversion results are presented. Various atmospheric models representative of polar stratospheric aerosols are used in the SAM-2 and correlative sensor analyses.
An ozonesonde launched near electrically active convection in Houston, TX on 5 September 2013 during the NASA DISCOVER-AQ project measured a large enhancement of ozone throughout the troposphere. A separate ozonesonde was launched from Smith Point, TX (approx. 58 km southeast of the Houston site) at approximately the same time as the launch from Houston and did not measure that enhancement. Furthermore, ozone profiles for the descent of both sondes agreed well with the ascending Smith Point profile, suggesting a highly localized event in both space and time in which an anomalously large enhancement of 70 - 100 ppbv appeared in the ascending Houston ozonesonde data. Compared to literature values, such an enhancement appears to be the largest observed to date. Potential sources of the localized ozone enhancement such as entrainment of urban or biomass burning emissions, downward transport from the stratosphere, photochemical production from lightning NO(sub x), and direct ozone production from corona discharges were investigated using model simulations. We conclude that the most likely explanation for the large ozone enhancement is direct ozone production by corona discharges. Integrating the enhancement seen in the Houston ozone profile and using the number of electrical discharges detected by the NLDN (or HLMA), we estimate a production of 2.48 x 10(exp. 28) molecules of ozone per flash which falls within the range of previously recorded values (9.89 x 10(exp. 26) - 9.82 x 10)exp. 28) molecules of ozone per flash). Since there is currently no parameterization for the direct production of ozone from corona discharges we propose the implementation of an equation into a chemical transport model. Ultimately, additional work is needed to further understand the occurrence and impact of corona discharges on tropospheric chemistry on short and long timescales.
Sounding rockets are suborbital launch vehicles capable of carrying scientific payloads to several hundred miles in altitude. These missions return a variety of scientific data including: chemical makeup and physical processes taking place in the atmosphere, natural radiation surrounding the Earth, data on the Sun, stars, galaxies and many other phenomena. In addition, sounding rockets provide a reasonably economical means of conducting engineering tests for instruments and devices to be used on satellites and other spacecraft prior to their use in these more expensive missions. Typically around thirty of these rockets are launched each year, from established ranges at Wallops Island, Virginia; Poker Flat Research Range, Alaska; White Sands Missile Range, New Mexico and from a number of ranges outside the United States. Many times launches are conducted from temporary launch ranges in remote parts of the world requiring considerable expense to transport and operate tracking radars. In order to support these missions, an inverse differential GPS system has been developed. The flight system consists of a small, inexpensive receiver, a preamplifier and a wrap-around antenna. A rugged, compact, portable ground station extracts GPS data from the raw payload telemetry stream, performs a real time differential solution and graphically displays the rocket's path relative to a predicted trajectory plot. In addition to generating a real time navigation solution, the system has been used for payload recovery, timing, data timetagging, precise tracking of multiple payloads and slaving of optical tracking systems for over the horizon acquisition. This paper discusses, in detail, the flight and ground hardware, as well as data processing and operational aspects of the system, and provides evidence of the system accuracy.