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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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NASA NTRS1991

Space transfer vehicle concepts and requirements study. Volume 1: Executive summary

A description of the study in terms of background, objectives, and issues is provided. NASA is currently studying new initiatives of space exploration involving both piloted and unpiloted missions to destinations throughout the solar system. Many of these missions require substantial improvements in launch vehicle and upper stage capabilities. This study provides a focused examination of the Space Transfer Vehicles (STV) required to perform these missions using the emerging national launch vehicle definition, the Space Station Freedom (SSF) definition, and the latest mission scenario requirements. The study objectives are to define preferred STV concepts capable of accommodating future exploration missions in a cost-effective manner, determine the technology development (if any) required to perform these missions, and develop a decision database of various programmatic approaches for the development of the STV family of vehicles. Special emphasis was given to examining space basing (stationing reusable vehicles at a space station), examining the piloted lunar mission as a primary design mission, and restricting trade studies to the high-performance, near-term cryogenics (LO2/LH2) as vehicle propellant. The study progressed through three distinct 6-month phases. The first phase concentrated on supporting a NASA 3 month definition of exploration requirements (the '90-day study') and during this phase developed and optimized the space-based point-of-departure (POD) 2.5-stage lunar vehicle. The second phase developed a broad decision database of 95 different vehicle options and transportation architectures. The final phase chose the three most cost-effective architectures and developed point designs to carry to the end of the study. These reference vehicle designs are mutually exclusive and correspond to different national choices about launch vehicles and in-space reusability. There is, however, potential for evolution between concepts.

Weber, Gary A.↗
NASA NTRS1989

Real-time approximate optimal guidance laws for the advanced launch system

An approach to optimal ascent guidance for a launch vehicle is developed using an expansion technique. The problem is to maximize the payload put into orbit subject to the equations of motion of a rocket over a rotating spherical earth. It is assumed that the thrust and gravitational forces dominate over the aerodynamic forces. It is shown that these forces can be separated by a small parameter epsilon, where epsilon is the ratio of the atmospheric scale height to the radius of the earth. The Hamilton-Jacobi-Bellman or dynamic programming equation is expanded in a series where the zeroth-order term (epsilon = 0) can be obtained in closed form. The zeroth-order problem is that of putting maximum payload into orbit subject to the equations of motion of a rocket in a vacuum over a flat earth. The neglected inertial and aerodynamic terms are included in higher order terms of the expansion, which are determined from the solution of first-order linear partial differential equations requiring only quadrature integrations. These quadrature integrations can be performed rapidly, so that real-time approximate optimization can be used to construct the launch guidance law.

Speyer, Jason L.↗
NASA NTRS2017

Recent Advancements in the Infrared Flow Visualization System for the NASA Ames Unitary Plan Wind Tunnels

The following details recent efforts undertaken at the NASA Ames Unitary Plan wind tunnels to design and deploy an advanced, production-level infrared (IR) flow visualization data system. Highly sensitive IR cameras, coupled with in-line image processing, have enabled the visualization of wind tunnel model surface flow features as they develop in real-time. Boundary layer transition, shock impingement, junction flow, vortex dynamics, and buffet are routinely observed in both transonic and supersonic flow regimes all without the need of dedicated ramps in test section total temperature. Successful measurements have been performed on wing-body sting mounted test articles, semi-span floor mounted aircraft models, and sting mounted launch vehicle configurations. The unique requirements of imaging in production wind tunnel testing has led to advancements in the deployment of advanced IR cameras in a harsh test environment, robust data acquisition storage and workflow, real-time image processing algorithms, and evaluation of optimal surface treatments. The addition of a multi-camera IR flow visualization data system to the Ames UPWT has demonstrated itself to be a valuable analyses tool in the study of new and old aircraft/launch vehicle aerodynamics and has provided new insight for the evaluation of computational techniques.

Infrared Thermography↗
NASA NTRS1993

Post-Optimality Analysis In Aerospace Vehicle Design

This analysis pertains to the applicability of optimal sensitivity information to aerospace vehicle design. An optimal sensitivity (or post-optimality) analysis refers to computations performed once the initial optimization problem is solved. These computations may be used to characterize the design space about the present solution and infer changes in this solution as a result of constraint or parameter variations, without reoptimizing the entire system. The present analysis demonstrates that post-optimality information generated through first-order computations can be used to accurately predict the effect of constraint and parameter perturbations on the optimal solution. This assessment is based on the solution of an aircraft design problem in which the post-optimality estimates are shown to be within a few percent of the true solution over the practical range of constraint and parameter variations. Through solution of a reusable, single-stage-to-orbit, launch vehicle design problem, this optimal sensitivity information is also shown to improve the efficiency of the design process, For a hierarchically decomposed problem, this computational efficiency is realized by estimating the main-problem objective gradient through optimal sep&ivity calculations, By reducing the need for finite differentiation of a re-optimized subproblem, a significant decrease in the number of objective function evaluations required to reach the optimal solution is obtained.

Braun, Robert D.↗
NASA NTRS2016

Application of Design of Experiments and Surrogate Modeling within the NASA Advanced Concepts Office, Earth-to-Orbit Design Process

Decisions made during early conceptual design have a large impact upon the expected life-cycle cost (LCC) of a new program. It is widely accepted that up to 80% of such cost is committed during these early design phases [1]. Therefore, to help minimize LCC, decisions made during conceptual design must be based upon as much information as possible. To aid in the decision making for new launch vehicle programs, the Advanced Concepts Office (ACO) at NASA Marshall Space Flight Center (MSFC) provides rapid turnaround pre-phase A and phase A concept definition studies. The ACO team utilizes a proven set of tools to provide customers with a full vehicle mass breakdown to tertiary subsystems, preliminary structural sizing based upon worst-case flight loads, and trajectory optimization to quantify integrated vehicle performance for a given mission [2]. Although the team provides rapid turnaround for single vehicle concepts, the scope of the trade space can be limited due to analyst availability and the manpower requirements for manual execution of the analysis tools. In order to enable exploration of a broader design space, the ACO team has implemented an advanced design methods (ADM) based approach. This approach applies the concepts of design of experiments (DOE) and surrogate modeling to more exhaustively explore the trade space and provide the customer with additional design information to inform decision making. This paper will first discuss the automation of the ACO tool set, which represents a majority of the development effort. In order to fit a surrogate model within tolerable error bounds a number of DOE cases are needed. This number will scale with the number of variable parameters desired and the complexity of the system's response to those variables. For all but the smallest design spaces, the number of cases required cannot be produced within an acceptable timeframe using a manual process. Therefore, automation of the tools was a key enabler for the successful application of an ADM approach to an ACO design study. Following the overview of the tool set automation, an example problem will be given to illustrate the implementation of the ADM approach. The example problem will first cover the inclusion of ground rules and assumptions (GR&A) for a study. The GR&A are very important to the study as they determine the constraints within which a trade study can be conducted. These trades must ultimately reconcile with the customer's desired output and any anticipated "what if" questions.

Zwack, Mathew R.↗
NASA NTRS2019

Small Launch Vehicle Sizing Analysis with Solid Rocket Examples

Small launch vehicles are governed by the same physics as large launch vehicles of course, but due to their small size, some aspects and sensitivities become more important and others less. This paper shows semi-empirical correlations to quantify dry mass fraction for both stage and whole vehicle optimization: mass fraction due to density, mass fraction due to thrust-to-weight, and mass fraction due to size reduction. For single-stage optimizations, a stage performance requirement can be met by a locus of mass fraction vs. specific impulse. Based on the above correlations, this alone can recommend a solid or liquid rocket for a stage. Rocket designs of similar technology levels are compared, focusing on where stages become less mass-efficient as they get smaller. The Mars Ascent Vehicle is shown to exemplify a trade between a two-stage solids vehicle and a one- or two-stage liquids vehicle.

Kibbey, Tim↗
NASA NTRS2019

Small Launch Vehicle Sizing Analysis With Solid Rocket Examples

Small launch vehicles are governed by the same physics as large launch vehicles of course, but due to their small size, some aspects and sensitivities become more important and others less. This paper shows semi-empirical correlations to quantify dry mass fraction for both stage and whole vehicle optimization: mass fraction due to density, mass fraction due to thrust-to-weight, and mass fraction due to size reduction. For single-stage optimizations, a stage performance requirement can be met by a locus of mass fraction vs. specific impulse. Based on the above correlations, this alone can recommend a solid or liquid rocket for a stage. Rocket designs of similar technology levels are compared, focusing on where stages become less mass-efficient as they get smaller. The Mars Ascent Vehicle is shown to exemplify a trade between a two-stage solids vehicle and a one- or two-stage liquids vehicle.

Kibbey, Tim↗
NASA NTRS2001

Performance Evaluation of the NASA GTX RBCC Flowpath

The NASA Glenn Research Center serves as NASAs lead center for aeropropulsion. Several programs are underway to explore revolutionary airbreathing propulsion systems in response to the challenge of reducing the cost of space transportation. Concepts being investigated include rocket-based combined cycle (RBCC), pulse detonation wave, and turbine-based combined cycle (TBCC) engines. The GTX concept is a vertical launched, horizontal landing, single stage to orbit (SSTO) vehicle utilizing RBCC engines. The propulsion pod has a nearly half-axisymmetric flowpath that incorporates a rocket and ram-scramjet. The engine system operates from lift-off up to above Mach 10, at which point the airbreathing engine flowpath is closed off, and the rocket alone powers the vehicle to orbit. The paper presents an overview of the research efforts supporting the development of this RBCC propulsion system. The experimental efforts of this program consist of a series of test rigs. Each rig is focused on development and optimization of the flowpath over a specific operating mode of the engine. These rigs collectively establish propulsion system performance over all modes of operation, therefore, covering the entire speed range. Computational Fluid Mechanics (CFD) analysis is an important element of the GTX propulsion system development and validation. These efforts guide experiments and flowpath design, provide insight into experimental data, and extend results to conditions and scales not achievable in ground test facilities. Some examples of important CFD results are presented.

Thomas, Scott R.↗
NASA NTRS1974

Program manual for HILTOP, a heliocentric interplanetary low thrust trajectory optimization program. Part 1: User's guide

A performance-analysis computer program, that was developed explicitly to generate optimum electric propulsion trajectory data for missions of interest in the exploration of the solar system is presented. The program was primarily designed to evaluate the performance capabilities of electric propulsion systems, and in the simulation of a wide variety of interplanetary missions. A numerical integration of the two-body, three-dimensional equations of motion and the Euler-Lagrange equations was used in the program. Transversality conditions which permit the rapid generation of converged maximum-payload trajectory data, and the optimization of numerous other performance indices for which no transversality conditions exist are included. The ability to simulate constrained optimum solutions, including trajectories having specified propulsion time and constant thrust cone angle, is also in the program. The program was designed to handle multiple-target missions with various types of encounters, such as rendezvous, stopover, orbital capture, and flyby. Performance requirements for a variety of launch vehicles can be determined.

Mann, F. I.↗
NASA NTRS1991

Monte Carlo simulation of launchsite winds at Kennedy Space Center

This paper develops and validates an easily implemented model for simulating random horizontal wind profiles over the Kennedy Space Center (KSC) at Cape Canaveral, Florida. The model is intended for use in Monte Carlo launch vehicle simulations of the type employed in mission planning, where the large number of profiles needed for statistical fidelity of such simulation experiments makes the use of actual wind measurements impractical. The model is based on measurements made at KSC and represents vertical correlations by a decaying exponential model which is parameterized via least-squares parameter optimization against the sample data. The validity of the model is evaluated by comparing two Monte Carlo simulations of an asymmetric, heavy-lift launch vehicle. In the first simulation, the measured wind profiles are used, while in the second, the wind profiles are generated using the stochastic model. The simulations indicate that the use of either the measured or simulated wind field results in similar launch vehicle performance.

Queen, Eric M.↗
NASA NTRS1993

Rocket-powered single-stage vehicle configuration selection and design

A reusable rocket-powered, single-stage launch vehicle has been designed as a part of NASA's Advanced Manned Launch System (AMLS) study to examine options for a next-generation manned space transportation system. The configuration selection process utilized a response surface methodology for multidisciplinary optimization. The methodology was utilized to determine the minimum dry weight entry vehicle to meet constraints on landing velocity and on subsonic, supersonic, and hypersonic trim and stability. Once the optimum configuration was determined, a multidisciplinary conceptual vehicle design was performed. This paper presents the results of the configuration selection methodology and summarizes the overall conceptual design process with special attention given to the individual disciplines of weights/ sizing, structures/materials, configuration, flight mechanics, aerodynamics, aeroheating, propulsion, and operations.

Stanley, Douglas O.↗
NASA NTRS2010

Modular Approach to Launch Vehicle Design Based on a Common Core Element

With a heavy lift launch vehicle as the centerpiece of our nation's next exploration architecture's infrastructure, the Advanced Concepts Office at NASA's Marshall Space Flight Center initiated a study to examine the utilization of elements derived from a heavy lift launch vehicle for other potential launch vehicle applications. The premise of this study is to take a vehicle concept, which has been optimized for Lunar Exploration, and utilize the core stage with other existing or near existing stages and boosters to determine lift capabilities for alternative missions. This approach not only yields a vehicle matrix with a wide array of capabilities, but also produces an evolutionary pathway to a vehicle family based on a minimum development and production cost approach to a launch vehicle system architecture, instead of a purely performance driven approach. The upper stages and solid rocket booster selected for this study were chosen to reflect a cross-section of: modified existing assets in the form of a modified Delta IV upper stage and Castor-type boosters; potential near term launch vehicle component designs including an Ares I upper stage and 5-segment boosters; and longer lead vehicle components such as a Shuttle External Tank diameter upper stage. The results of this approach to a modular launch system are given in this paper.

Creech, Dennis M.↗
NASA NTRS2015

Suborbital Asteroid Intercept and Fragmentation for Very Short Warning Time Scenarios

Small near-Earth objects (NEOs) 50150 m in size are far more numerous (hundreds of thousands to millions yet to be discovered) than larger NEOs. Small NEOs, which are mostly asteroids rather than comets, are very faint in the night sky due to their small sizes, and are, therefore, difficult to discover far in advance of Earth impact. However, even small NEOs are capable of creating explosions with energies on the order of tens or hundreds of megatons (Mt).We are, therefore, motivated to prepare to respond effectively to short warning time, small NEO impact scenarios. In this paper we explore the lower bound on actionable warning time by investigating the performance of notional upgraded Intercontinental Ballistic Missiles (ICBMs) to carry Nuclear Explosive Device (NED) payloads to intercept and disrupt a fictitious incoming NEO at high altitudes (generally, at least 2500 km above Earth). We conduct this investigation by developing optimal NEO intercept trajectories for a range of cases and comparing their performances.Our results show that suborbital NEO intercepts using Minuteman III or SM-3 IIA launch vehicles could achieve NEO intercept a few minutes prior to when the NEOwould strike Earth. We also find that more powerful versions of the launch vehicles (e.g., total V 9.511 kms) could intercept incoming NEOs over a day prior to when the NEO would strike Earth, if launched at least several days prior to the time of NEO intercept. Finally, we discuss a number of limiting factors and practicalities that affect whether the notional systems we describe could become feasible.

Planetary Defense↗
NASA NTRS1998

Subsystem Analysis/Optimization for the X-34 Main Propulsion System

The Orbital Sciences Corporation X-34 vehicle demonstrates technologies and operations key to future reusable launch vehicles. The general flight performance goal of this unmanned rocket plane is Mach 8 flight at an altitude of 250,000 feet. The Main Propulsion System (MPS) supplies liquid propellants to the main engine, which provides the primary thrust for attaining mission goals. Major MPS design and operational goals are aircraft-like ground operations, quick turnaround between missions, and low initial/operational costs. Analyses related to optimal MPS subsystem design are reviewed in this paper. A pressurization system trade weighs maintenance/reliability concerns against those for safety in a comparison of designs using pressure regulators versus orifices to control pressurant flow. A propellant dump/feed system analysis weighs the issues of maximum allowable vehicle landing weight, trajectory, and MPS complexity to arrive at a final configuration for propellant dump/feed systems.

McDonald, J. P.↗
NASA NTRS2011

Launch Vehicle Demonstrator Using Shuttle Assets

The Advanced Concepts Office at NASA's George C. Marshall Space Flight Center undertook a study to define candidate early heavy lift demonstration launch vehicle concepts derived from existing space shuttle assets. The objective was to determine the performance capabilities of these vehicles and characterize potential early demonstration test flights. Given the anticipated budgetary constraints that may affect America's civil space program, and a lapse in U.S. heavy launch capability with the retirement of the space shuttle, an early heavy lift launch vehicle demonstration flight would not only demonstrate capabilities that could be utilized for future space exploration missions, but also serve as a building block for the development of our nation s next heavy lift launch system. An early heavy lift demonstration could be utilized as a test platform, demonstrating capabilities of future space exploration systems such as the Multi Purpose Crew Vehicle. By using existing shuttle assets, including the RS-25D engine inventory, the shuttle equipment manufacturing and tooling base, and the segmented solid rocket booster industry, a demonstrator concept could expedite the design-to-flight schedule while retaining critical human skills and capital. In this study two types of vehicle designs are examined. The first utilizes a high margin/safety factor battleship structural design in order to minimize development time as well as monetary investment. Structural design optimization is performed on the second, as if an operational vehicle. Results indicate low earth orbit payload capability is more than sufficient to support various vehicle and vehicle systems test programs including Multi-Purpose Crew Vehicle articles. Furthermore, a shuttle-derived, hydrogen core vehicle configuration offers performance benefits when trading evolutionary paths to maximum capability.

Creech, Dennis M.↗
NASA NTRS1977

Trajectory optimization for the Atlas/Centaur launch vehicle

A method for finding solutions to complex trajectory optimization problems with detailed hardware and launch vehicle constraints is presented. The trajectory optimization problem is formulated as a constrained function minimization problem by modeling the control by a function of a finite number of parameters. Although any parametric control model may be selected, the natural choice of parameters is the actual parameters of the parametric guidance equations flown on the flight digital computer. A variation of the method of multipliers is presented for performing the constrained function minimization. A functional is proposed that allows incorporation of state variable inequality constraints. Finally, optimization of the HEAO-A direct-ascent trajectory is discussed. The HEAO-A mission demonstrates profile optimization to meet range safety, tracking, heating, structural integrity, and attitude rate constraints.

Brusch, R. G.↗
NASA NTRS1996

Optimization Methodology for Unconventional Rocket Nozzle Design

Several current rocket engine concepts such as the bell-annular tripropellant engine, and the linear aerospike being proposed for the X-33, require unconventional three-dimensional rocket nozzles which must conform to rectangular or sector-shaped envelopes to meet integration constraints. These types of nozzles exist outside the current experience database, therefore, development of efficient design methods for these propulsion concepts is critical to the success of launch vehicle programs. Several approaches for optimizing rocket nozzles, including streamline tracing techniques, and the coupling of CFD analysis to optimization algorithms are described. The relative strengths and weaknesses of four classes of optimization algorithms are discussed: Gradient based methods, genetic algorithms, simplex methods, and surface response methods. Additionally, a streamline tracing technique, which provides a very computationally efficient means of defining a three-dimensional contour, is discussed. The performance of the various optimization methods on thrust optimization problems for tripropellant and aerospike concepts is assessed and recommendations are made for future development efforts.

Follett, W.↗
NASA NTRS1982

Black Brant X third stage dispersion dynamics

The successful development of the Black Brant X (BBX) launch vehicle resulted from the application of a gyroscopically stabilized, unguided third stage motor configuration. Design analyses indicated sufficient gyroscopic stability and acceptable impact dispersion could be realized by optimizing second/third stage separation time. By flying an unguided third stage motor, significant program cost savings were realized. One flight test and two operational missions demonstrated the feasibility of utilizing gyroscopic stability to achieve satisfactory trajectory dispersion on a high performance rocket. A detailed assessment of the major dispersion contributors identified for the first three BBX missions is presented, offering important insight into the dispersion sensitivity of unguided, spin-stabilized flight.

Montag, W. H.↗