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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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An evaluation is presented of J-2 engine modifications that will simplify operation and improve reliability of the advanced Saturn 1C launch vehicle. Methods of increasing thrust without extensively modifying the S-2 or S-4B stages are also evaluated. A thrust increase was achieved by raising engine combustion through a redesign of the engine thrust chamber and propellant feed system.
Flight-mechanics problems associated with large-scale transport of lunar mass to a space colony or manufacturing facility are discussed. The proposed transport method involves launch of payloads from a mass-driver on the lunar surface, onto ballistic trajectories to a passive mass-catcher located near the L2 libration point, with the caught mass subsequently being transported to the colony. Arrival velocities at L2, sensitivities in arrival dispersion due to launch errors, and effects of launch site location are treated, via numerically integrated orbits in the restricted three-body problem. From any launch site it is possible to define a target point reached with zero dispersion due to errors in a selected component of launch velocity. Effects of lunar geometrical librations and of obliquity, as well as the conditions for biasing a trajectory away from L2 so as to reduce stationkeeping costs, are dealt with along with transfer orbits from L2 to the colony. The theory of capture and the theory of resonance lead to a colony orbit, with period approximately two weeks, reached from L2 with velocity increment as low as 9.02 m/sec.
A summary is given of the concept definition phase of NASA's Orbital Transfer Vehicle (OTV) program. The mission requirements are threefold: (1) the ability to deliver payloads well beyond the capability of any current system at low acceleration levels, (2) the extension of the cost effectiveness from reusability and operational flexibility provided by the Shuttle at LEO to the geosynchronous altitude, and (3) the provision of the round trip capability needed to service and maintain this expanded and more valuable resource at GEO. The initial Shuttle-matched, reusable OTV is shown to be cost effective with low development cost, relatively short development period and high projected reliability. The performance enhancement provided by aero-assist is discussed, which will enable the difficult round trip to be flown in a single Shuttle launch. Two methods of reducing vehicle velocity are available: (1) aeromaneuvering, which uses vehicle lift to climb or descend and thereby correct for density variations, and (2) aerobraking which varies vehicle drag directly to correct for density.
Due to large amounts of oxygen required for space travel, a method of mining, transporting, and storing this oxygen in space would facilitate further space exploration. The following project deals specifically with the methods for transporting liquid oxygen from the lunar surface to the Lunar Orbit (LO) space station, and then to the Lower Earth Orbit (LEO) space station. Two vehicles were designed for operation between the LEO and LO space stations. The first of these vehicles is an aerobraked design vehicle. The Aerobrake Orbital Transfer Vehicle (OTV) is capable of transporting 5000 lbm of payload to LO while returning to LEO with 60,000 lbm of liquid oxygen, and thus meet mission requirements. The second vehicle can deliver 18,000 lbm of payload to LO and is capable of bringing 60,000 lbm of liquid oxygen back to LEO. A lunar landing vehicle was also designed for operation between LO and the established moon base. The use of an electromagnetic railgun as a method for launching the lunar lander was also investigated. The feasibility of the railgun is doubtful at this time. A system of spheres was also designed for proper storing and transporting of the liquid oxygen. The system assumes a safe means for transferring the liquid oxygen from tank to tank is operational. A sophisticated life support system was developed for both the OTV and the lunar lander. This system focuses on such factors as the vehicle environment, waste management, water requirements, food requirements, and oxygen requirements.
A winch system provides a method for launch and recovery capabilities for kites and tethered blimps or balloons. Low power consumption is a key objective, as well as low weight for portability. This is accomplished by decoupling the tether-line storage and wind ing/ unwinding functions, and providing tailored and efficient mechanisms for each. The components of this system include rotational power input devices such as electric motors or other apparatus, line winding/unwinding reel(s), line storage reel(s), and independent drive trains. Power is applied to the wind/unwind reels to transport the tether line. Power is also applied to a line storage reel, from either the wind/unwind power source, the wind/unwind reel itself, or separate power source. The speeds of the two reels are synchronized, but not dependent on each other. This is accomplished via clutch mechanisms, variable transmissions, or independent motor controls. The speed of the storage reel is modulated as the effective diameter of the reel changes with line accumulation.
Lasers with high peak power pulses are commonly used in spaceflight missions for a wide range of applications, from LIDAR systems to optical communications. Due to the high optical power needed, the laser has to be located on the exterior of the satellite or coupled through a series of free space optics. This presents challenges for thermal management, radiation resistance, and mechanical design. Future applications will require multiple lasers located close together, which further complicates the design. Coupling the laser energy into a fiber optic cable allows the laser to be relocated to a more favorable position on the spacecraft. Typical fiber optic termination procedures are not sufficient for injection of these high-power laser pulses without catastrophic damage to the fiber endface. In the current study, we will review the causes of fiber damage during high-power injection and discuss our new manufacturing procedures that overcome these issues to permit fiber use with high reliability in these applications. We will also discuss the proper methods for launching the laser pulses into the fiber to avoid damage and how this is being implemented for current spaceflight missions.
The calibration of the historical AVHRR visible channels has always been hindered by the degrading NOAA satellite orbits. The complete AVHRR record spans over three decades making it useful for cloud, aerosol, and land use climate studies. Studies that monitor long-term changes in these properties require climate quality calibration. The AVHRR sensors do not have any onboard visible calibration, unlike the IR channels. Many AVHRR post-launch calibration methods have been published. Some of these are based on the pseudo invariant targets such as desert and polar ice and others employ direct calibration transfer from a well-calibrated sensor such as MODIS. Most historical studies only employed only one method and reconciling calibration differences from multiple studies was difficult. This study employs multiple calibration approaches, which are then merged according to their individual uncertainty. Since the visible spectral response functions are similar across AVHRR sensor and that the morning and afternoon orbits usually degrade in the same manner, an afternoon and morning orbiter during the MODIS era is chosen as the reference AVHRR sensor. The reference AVHRR sensor is inter-calibrated with the Aqua-MODIS Collection 6 standard using simultaneous nadir overpass (SNO) radiance pairs. Three pseudo invariant target approaches are utilized, deserts, polar ice and deep convective clouds. These sites are characterized using reference AVHRR reflectances as a function of solar zenith angle. The multiple calibration approach is then validated with non-reference AVHRR sensor during the MODIS era, using AVHRR and MODIS SNO measurements. For historical AVHRR satellite having consistent calibration across invariant targets then validates the approach. This calibration effort is in sponsored by the NOAA CDR program to provide climate quality visible calibration coefficients to the remote sensing community. Each of the calibration approaches are unique and when utilized in tandem can accurately calibrate the AVHRR record. AVHRR calibration methods cannot rely on repeatable annual orbits for calibration model development. Inconsistent results prompt code and invariant target examinations that cannot be determined using only single method approaches. Results from this study will be presented at the meeting.
A method and an associated digital computer program for evaluating the vibrational characteristics of large liquid-filled rigid wall tanks of general shape are presented. A solution procedure was developed in which slosh modes and frequencies are computed for systems mathematically modeled as assemblages of liquid finite elements. To retain sparsity in the assembled system mass and stiffness matrices, a compressible liquid element formulation was incorporated in the program. The approach taken in the liquid finite element formulation is compatible with triangular and quadrilateral structural finite elements so that the analysis of liquid motion can be coupled with flexible tank wall motion at some future time. The liquid element repertoire developed during the course of this study consists of a two-dimensional triangular element and a three-dimensional tetrahedral element.
A method for accurately initializing spacecraft attitude after release from the Orbiter is described. It is noted that the method is suitable for an autonomous ascent to mission orbit. Test results are given from a FORTRAN simulation of the estimation algorithm using measurement data from a detailed spacecraft dynamics simulation program. The technique here is orbital yaw-gyrocompassing. Attitude is estimated through a Kalman filter, using pitch and roll measurements from an earth sensor, while gyro data provide the system dynamics information. In the tests described, gyro and earth sensor data are generated by an existing control system simulation of earth-search and yaw-gyrocompassing attitude dynamics; they include realistic errors such as delays, random noise and quantization effects. The estimated attitude history is compared with the true attitude history from the simulation program to assess the accuracy and convergence of the filter in the presence of noisy measurements and disturbances, including thruster firings for momentum control. It is noted that since the earth sensor provides direct measurements of pitch and roll, the main criterion of filter performance is yaw accuracy.
A method is proposed that provides advance information about unpredictable atmospheric density dispersions that must be accommodated during random operations of aeroassisted-orbital-transfer vehicles (AOTVs). The principal feature is that a test or 'scout' projectile precedes the AOTV through the same region of the atmosphere as that of the predicted transatmospheric flight trajectory. The atmospheric density structure is determined from the vehicle's aerodynamic deceleration characteristics by on-board or ground-based tracking equipment. The time lag between passage of the projectile and the AOTV can be adjusted to only that time necessary to implement required guidance, navigation, and control (GN&C) corrections. The various strategies available to control the projectile's flight characteristics are analyzed in detail. The results are correlated with aerothermodynamic heating and materials requirements to ensure the survival of the projectile and, consequently, the capability of the AOTV to navigate a variable upper atmosphere within specified limits.
Advanced launch systems (e.g., Reusable Launch Vehicle and other Shuttle Class concepts, Rocket-Based Combine Cycle, etc.), and interplanetary vehicles will very likely incorporate fiber reinforced ceramic matrix composites (CMC) in critical propulsion components. The use of CMC is highly desirable to save weight, to improve reuse capability, and to increase performance. CMC candidate applications are mission and cycle dependent and may include turbopump rotors, housings, combustors, nozzle injectors, exit cones or ramps, and throats. For reusable and single mission uses, accurate prediction of life is critical to mission success. The tools to accomplish life prediction are very immature and not oriented toward the behavior of carbon fiber reinforced silicon carbide (C/SiC), the primary system of interest for a variety of space propulsion applications. This paper describes an approach to satisfy the need to develop an integrated life prediction system for CMC that addresses mechanical durability due to cyclic and steady thermomechanical loads, and takes into account the impact of environmental degradation.
This paper provides an overview of the U.S. space nuclear power system launch approval process as defined by the two separate requirements of the National Environmental Policy Act (NEPA) and Presidential Directive/National Security Council Memorandum No. 25 (PD/NSC-25).
This slide presentation reviews the use of terahertz imaging and Backscatter Radiography in a probability of detection study of the foam on the external tank (ET) shedding and damaging the shuttle orbiter. Non-destructive Examination (NDE) is performed as one method of preventing critical foam debris during the launch. Conventional NDE methods for inspection of the foam are assessed and the deficiencies are reviewed. Two methods for NDE inspection are reviewed: Backscatter Radiography (BSX) and Terahertz (THZ) Imaging. The purpose of the Probability of Detection (POD) study was to assess performance and reliability of the use of BSX and or THZ as an appropriate NDE method. The study used a test article with inserted defects, and a sample of blanks included to test for false positives. The results of the POD study are reported.
A system, method, and apparatus for a vented launch vehicle adaptor (LVA) for a manned spacecraft with a "pusher" launch abort system are disclosed. The disclosed LVA provides a structural interface between a commercial crew vehicle (CCV) crew module/service module (CM/SM) spacecraft and an expendable launch vehicle. The LVA provides structural attachment of the module to the launch vehicle. It also provides a means to control the exhaust plume from a pusher-type launch abort system that is integrated into the module. In case of an on-pad or ascent abort, which requires the module to jettison away from the launch vehicle, the launch abort system exhaust plume must be safely directed away from critical and dangerous portions of the launch vehicle in order to achieve a safe and successful jettison.
Norton-Thevenin Receptance Coupling (NTRC), as described in several NASA Engineering and Safety Center (NESC) papers shows promise in enabling loads development for payloads with less computational cost and analyst time as compared to a full integrated vehicle coupled loads analysis (CLA). NTRC allows the free vehicle responses and impedance (accelerance) at the payload to vehicle interface (derived from integrated vehicle CLA) to be used with a payload model in a payload CLA. Due to the NTRC damping differing from the full integrated vehicle damping, results are slightly different, so coverage factors were developed to ensure NTRC results enveloped results from a traditional CLA. NTRC results with coverage factors were very close to full CLA results. The Space Launch System (SLS) coupled loads team has developed an implementation of the NTRC method to enable support of co-manifested payloads with analysis needs that do not fit the primary SLS load cycle schedule. The NTRC method was successfully used to support the European System Providing Refueling, Infrastructure, and Communications (ESPRIT) module for Gateway planned to fly on Artemis V and has been adopted as the SLS approach for supporting payload CLA requests that do not align with vehicle load cycles.
Numerical methods and computer programs for studying the stability and evolution of orbits of large eccentricity are presented. Methods for determining launch windows and target dates are developed. Mathematical models are prepared to analyze the characteristics of specific missions.
Probability of favorable weather for successful aerospace vehicle launch with known operational launch contraints, using empirical method of determination
The purpose of this thesis is to design a machine learning algorithm platform that provides expanded knowledge of mission availability through a launch season by improving trajectory resolution and introducing launch mission forecasting. The specific scenario addressed in this paper is one in which data is provided for four deterministic translational maneuvers through a mission to a Near Rectilinear Halo Orbit (NRHO) with a 9:2 synodic frequency. Current launch availability knowledge under NASA's Orion Orbit Performance Team is established by altering optimization variables associated to given reference launch epochs. This current method can bean abstract task and relies on an orbit analyst to structure a mission based on an established mission design methodology associated to the performance of Orion and NASA's Space Launch System. Introducing a machine learning algorithm trained to construct mission scenarios within the feasible range of known trajectories reduces the required interaction of the orbit analyst by removing the needed step of optimizing the orbit to fit an expected translational response required of the spacecraft. In this study, k-Nearest Neighbor and Bayesian Linear Regression successfully predicted classical orbital elements for the launch windows observed. However both algorithms had limitations due to their approaches to model fitting. Training machine learning algorithms off of classical orbital elements introduced a repetitive approach to reconstructing mission segments for different arrival opportunities through the launch window and can prove to be a viable method of launch window scan generation for future missions.