SEARCH · Search NASA
Results for “launch vehicle performance”
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.
SPACECRAFT AND LAUNCH-VEHICLE PERFORMANCE
Spacecraft and launch vehicle performance - ma-7
Study of direct versus orbital entry for Mars missions. Volume 3 - Appendix A - Launch vehicle performance and flight mechanics
Titan 3 launch vehicle performance and flight mechanics analyses - direct versus orbital entry for Mars missions
Variations in launch vehicle performance from a mission planning viewpoint
Performance prediction for launch vehicle mission and payload determination
Laser Powered Launch Vehicle Performance Analyses
The purpose of this study is to establish the technical ground for modeling the physics of laser powered pulse detonation phenomenon. Laser powered propulsion systems involve complex fluid dynamics, thermodynamics and radiative transfer processes. Successful predictions of the performance of laser powered launch vehicle concepts depend on the sophisticate models that reflects the underlying flow physics including the laser ray tracing the focusing, inverse Bremsstrahlung (IB) effects, finite-rate air chemistry, thermal non-equilibrium, plasma radiation and detonation wave propagation, etc. The proposed work will extend the base-line numerical model to an efficient design analysis tool. The proposed model is suitable for 3-D analysis using parallel computing methods.
Launch Vehicle Performance for Bipropellant Propulsion Using Atomic Propellants With Oxygen
Atomic propellants for bipropellant launch vehicles using atomic boron, carbon, and hydrogen were analyzed. The gross liftoff weights (GLOW) and dry masses of the vehicles were estimated, and the 'best' design points for atomic propellants were identified. Engine performance was estimated for a wide range of oxidizer to fuel (O/F) ratios, atom loadings in the solid hydrogen particles, and amounts of helium carrier fluid. Rocket vehicle GLOW was minimized by operating at an O/F ratio of 1.0 to 3.0 for the atomic boron and carbon cases. For the atomic hydrogen cases, a minimum GLOW occurred when using the fuel as a monopropellant (O/F = 0.0). The atomic vehicle dry masses are also presented, and these data exhibit minimum values at the same or similar O/F ratios as those for the vehicle GLOW. A technology assessment of atomic propellants has shown that atomic boron and carbon rocket analyses are considered to be much more near term options than the atomic hydrogen rockets. The technology for storing atomic boron and carbon has shown significant progress, while atomic hydrogen is not able to be stored at the high densities needed for effective propulsion. The GLOW and dry mass data can be used to estimate the cost of future vehicles and their atomic propellant production facilities. The lower the propellant's mass, the lower the overall investment for the specially manufactured atomic propellants.
Launch Vehicle Performance with Solid Particle Feed Systems for Atomic Propellants
An analysis of launch vehicle Gross Liftoff Weight (GLOW) using high energy density atomic propellants with solid particle feed systems was conducted. The analyses covered several propellant combinations, including atoms of aluminum (Al), boron (B). carbon (C), and hydrogen (H) stored in a solid cryogenic particle, with a cryogenic liquid as the carrier fluid. Several different weight percents (wt%) for the liquid carrier were investigated and the gross lift off weight (GLOW) of the vehicles using the solid particle feed systems were compared with a conventional 02/H2 propellant vehicle. The potential benefits and effects of feed systems using solid particles in a liquid cryogenic fluid are discussed.
Launch vehicle performance using metallized propellants
Metallized propellant propulsion systems are considered as replacements for the solid rocket boosters and liquid sustainer stages on the current launch vehicles: both the Space Transportation System (STS) and the Titan 4. Liquid rocket boosters for the STS were analyzed as replacements for current solid rocket boosters. These boosters can provide a liquid propulsion system within the volume constraints of a solid rocket booster. A replacement for the Space Shuttle Main Engines using metallized O2/H2/Al was studied. The liquid stages of the Titan 4 were also investigated; the Aerozine-50 (A-50) fuel was replaced with metallized storable A-50/Al. A metallized propellant is similar to a traditional liquid propellant. However, it has metal particles, such as aluminum, that are suspended in a gelled fuel, such as hydrogen, RP-1, A-50 or monomethyl hydrazine (MMH). The fuels then undergo combustion with liquid oxygen or nitrogen tetroxide (NTO). These propellants provide options for increasing the performance of existing launch vehicle chemical propulsion systems by increasing fuel density or specific impulse or both. These increases in density and specific impulse can significantly reduce the propulsion system liftoff weight and allow a liquid rocket booster to fit into the same volume as an existing solid rocket booster. Also, because gelled fuels are akin to liquid propellants, metallized systems can provide enhanced controllability over solid propulsion systems. Gelling of the propellant also reduces the sensitivity to impacts and consequently reduces the propellant explosion hazard.
Launch vehicle performance using metallized propellants
Metallized propellant propulsion systems are considered as replacements for the solid rocket boosters and liquid sustainer stages on the current launch vehicles: both the Space Transportation System (STS) and the Titan 4. Liquid rocket boosters for the STS were analyzed as replacements for current solid rocket boosters. These boosters can provide a liquid propulsion system within the volume constraints of a solid rocket booster. A replacement for the Space Shuttle Main Engines using metallized 02/H2/A1 was studied. The liquid stages of the Titan 4 were also investigated; the Aerozine-50 (A-50) fuel was replaced with metallized storable A-50/A1. A metallized propellant is similar to a traditional liquid propellant. However, it has metal particles, such as aluminum, that are suspended in a gelled fuel, such as hydrogen, RP-1, A-50 or monomethyl hydrazine (MMH). The fuels then undergo combustion with liquid oxygen or nitrogen tetroxide (NTO). These propellants provide options for increasing the performance of existing launch vehicle chemical propulsion systems by increasing fuel density or specific impulse or both. These increases in density and specific impulse can significantly reduce the propulsion system liftoff weight and allow a liquid rocket booster to fit into the same volume as an existing solid rocket booster. Also, because gelled fuels are akin to liquid propellants, metallized systems can provide enhanced controllability over solid propulsion systems. Gelling of the propellant also reduces the sensitivity to impacts and consequently reduces the propellant explosion hazard.
Net Shape Spin Formed Cryogenic Aluminum Lithium Cryogenic Tank Domes for Lower Cost Higher Performance Launch Vehicles
With the goal of lower cost (simplified manufacturing and lower part count) and higher performance (higher strength to weight alloys) the NASA Technical Maturation Program in 2006 funded a proposal to investigate spin forming of space launch vehicle cryogenic tank domes. The project funding continued under the NASA Exploration Technology Development Program through completion in FY12. The first phase of the project involved spin forming of eight, 1 meter diameter "path finder" domes. Half of these were processed using a concave spin form process (MT Aerospace, Augsburg Germany) and the other half using a convex process (Spincraft, Boston MA). The convex process has been used to produce the Ares Common Bulkhead and the concave process has been used to produce dome caps for the Space Shuttle light weight external tank and domes for the NASDA H2. Aluminum Lithium material was chosen because of its higher strength to weight ratio than the Aluminum 2219 baseline. Aluminum lithium, in order to obtain the desired temper (T8), requires a cold stretch after the solution heat treatment and quench. This requirement favors the concave spin form process which was selected for scale up. This paper describes the results of processing four, 5.5 meter diameter (upper stage scale) net shaped spin formed Aluminum Lithium domes. In order to allow scalability beyond the limits of foundry and rolling mills (about 12 foot width) the circular blank contained one friction stir weld (heavy lifter scales require a flat blank containing two welds). Mechanical properties data (tensile, fracture toughness, stress corrosion, and simulated service testing) for the parent metal and weld will also be discussed.
AAP-1A launch vehicle performance and preliminary range safety analyses. Volume 2 - Range safety
Range safety data with trajectories, impact, and casualty probability for Saturn 1B/AAP-1A launch vehicle
Space shuttle launch vehicle performance trajectory, exchange ratios, and dispersion analysis
A baseline space shuttle performance trajectory for Mission 3A launched from WTR has been generated. Design constraints of maximum dynamic pressure, longitudinal acceleration, and delivered payload were satisfied. Payload exchange ratios are presented with explanation on use. Design envelopes of dynamic pressure, SRB staging point, aerodynamic heating and flight performance reserves are calculated and included.
Performance Characteristics of the Little Joe Launch Vehicle
Performance characteristics of Little Joe launch vehicle
Preliminary MIPCC Enhanced F-4 and F-15 Performance Characteristics for a First Stage Reusable Launch Vehicle
Performance increases in turbojet engines can theoretically be achieved through Mass Injection Pre-Compressor Cooling (MIPCC), a process involving injecting water or oxidizer or both into an afterburning turbojet engine. The injection of water results in pre-compressor cooling, allowing the propulsion system to operate at high altitudes and Mach numbers. In this way, a MIPCC-enhanced turbojet engine could be used to power the first stage of a reusable launch vehicle or be integrated into an existing aircraft that could launch a 100-lbm payload to a reference 100-nm altitude orbit at 28 deg inclination. The two possible candidates for MIPCC flight demonstration that are evaluated in this study are the F-4 Phantom II airplane and the F-15 Eagle airplane (both of McDonnell Douglas, now The Boeing Company, Chicago, Illinois), powered by two General Electric Company (Fairfield, Connecticut) J79 engines and two Pratt & Whitney (East Hartford, Connecticut) F100-PW-100 engines, respectively. This paper presents a conceptual discussion of the theoretical performance of each of these aircraft using MIPCC propulsion techniques. Trajectory studies were completed with the Optimal Trajectories by Implicit Simulation (OTIS) software (NASA Glenn Research Center, Cleveland, Ohio) for a standard F-4 airplane and a standard F-15 airplane. Standard aircraft simulation models were constructed, and the thrust in each was altered in accordance with estimated MIPCC performance characteristics. The MIPCC and production aircraft model results were then reviewed to assess the feasibility of a MIPCC-enhanced propulsion system for use as a first-stage reusable launch vehicle; it was determined that the MIPCC-enhanced F-15 model showed a significant performance advantage over the MIPCC-enhanced F-4 model.
Preliminary MIPCC Enhanced F-4 and F-15 Preformance Characteristics for a First Stage Reusable Launch Vehicle
Performance increases in turbojet engines can theoretically be achieved through Mass Injection Pre-Compressor Cooling (MIPCC), a process involving injecting water or oxidizer or both into an afterburning turbojet engine. The injection of water results in pre-compressor cooling, allowing the propulsion system to operate at high altitudes and Mach numbers. In this way, a MIPCC-enhanced turbojet engine could be used to power the first stage of a reusable launch vehicle or be integrated into an existing aircraft that could launch a 100-lbm payload to a reference 100-nm altitude orbit at 28 deg inclination. The two possible candidates for MIPCC flight demonstration that are evaluated in this study are the F-4 Phantom II airplane and the F-15 Eagle airplane (both of McDonnell Douglas, now The Boeing Company, Chicago, Illinois), powered by two General Electric Company (Fairfield, Connecticut) J79 engines and two Pratt & Whitney (East Hartford, Connecticut) F100-PW-100 engines, respectively. This paper presents a conceptual discussion of the theoretical performance of each of these aircraft using MIPCC propulsion techniques. Trajectory studies were completed with the Optimal Trajectories by Implicit Simulation (OTIS) software (NASA Glenn Research Center, Cleveland, Ohio) for a standard F-4 airplane and a standard F-15 airplane. Standard aircraft simulation models were constructed, and the thrust in each was altered in accordance with estimated MIPCC performance characteristics. The MIPCC and production aircraft model results were then reviewed to assess the feasibility of a MIPCC-enhanced propulsion system for use as a first-stage reusable launch vehicle; it was determined that the MIPCC-enhanced F-15 model showed a significant performance advantage over the MIPCC-enhanced F-4 model.
MERCURY-REDSTONE LAUNCH-VEHICLE DEVELOPMENT AND PERFORMANCE
Mercury-redstone launch vehicle development and performance
Performance Efficient Launch Vehicle Recovery and Reuse
For decades, economic reuse of launch vehicles has been an elusive goal. Recent attempts at demonstrating elements of launch vehicle recovery for reuse have invigorated a debate over the merits of different approaches. The parameter most often used to assess the cost of access to space is dollars-per-kilogram to orbit. When comparing reusable vs. expendable launch vehicles, that ratio has been shown to be most sensitive to the performance lost as a result of enabling the reusability. This paper will briefly review the historical background and results of recent attempts to recover launch vehicle assets for reuse. The business case for reuse will be reviewed, with emphasis on the performance expended to recover those assets, and the practicality of the most ambitious reuse concept, namely propulsive return to the launch site. In 2015, United Launch Alliance (ULA) announced its Sensible, Modular, Autonomous Return Technology (SMART) reuse plan for recovery of the booster module for its new Vulcan launch vehicle. That plan employs a non-propulsive approach where atmospheric entry, descent and landing (EDL) technologies are utilized. Elements of such a system have a wide variety of applications, from recovery of launch vehicle elements in suborbital trajectories all the way to human space exploration. This paper will include an update on ULA's booster module recovery approach, which relies on Hypersonic Inflatable Aerodynamic Decelerator (HIAD) and Mid-Air Retrieval (MAR) technologies, including its concept of operations (ConOps). The HIAD design, as well as parafoil staging and MAR concepts, will be discussed. Recent HIAD development activities and near term plans including scalability, next generation materials for the inflatable structure and heat shield, and gas generator inflation systems will be provided. MAR topics will include the ConOps for recovery, helicopter selection and staging, and the state of the art of parachute recovery systems using large parafoils for space asset recovery and high altitude deployment. The next proposed HIAD flight demonstration is called HULA (for HIAD on ULA), and will feature a 6m diameter HIAD. An update for the HULA concept will be provided in this paper. As proposed, this demonstration will fly as a secondary payload on an Atlas mission. The Centaur upper stage provides the reentry pointing, deorbit burn, and entry vehicle spin up. The flight test will culminate with a recovery of the HIAD using MAR. HULA will provide data from a Low Earth Orbit (LEO) return aeroheating environment that enables predictive model correlation and refinement. The resultant reduction in performance uncertainties should lead to design efficiencies that are increasingly significant at larger scales. Relevance to human scale Mars EDL using a HIAD will also be presented, and the applicability of the data generated from both HULA and SMART Vulcan flights, and its value for NASA's human exploration efforts will be discussed. A summary and conclusion will follow.