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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 NTRS2015

Pressure-Equalizing Cradle for Booster Rocket Mounting

A launch system and method improve the launch efficiency of a booster rocket and payload. A launch aircraft atop which the booster rocket is mounted in a cradle, is flown or towed to an elevation at which the booster rocket is released. The cradle provides for reduced structural requirements for the booster rocket by including a compressible layer, that may be provided by a plurality of gas or liquid-filled flexible chambers. The compressible layer contacts the booster rocket along most of the length of the booster rocket to distribute applied pressure, nearly eliminating bending loads. Distributing the pressure eliminates point loading conditions and bending moments that would otherwise be generated in the booster rocket structure during carrying. The chambers may be balloons distributed in rows and columns within the cradle or cylindrical chambers extending along a length of the cradle. The cradle may include a manifold communicating gas between chambers.

Rutan, Elbert L.↗
NASA NTRS2019

Aerodynamic Characterization and Improved Testing Methods for the Space Launch System Liftoff and Transition Environment

Low-speed wind tunnel testing for the liftoff and transition environment of the Space Launch System (SLS) was recently completed in the NASA Langley Research Center 14- by 22-Foot Subsonic Tunnel using 1.75%-scale models of three SLS vehicle configurations. During the liftoff testing, the primary objective was to evaluate the aerodynamic forces and moments on the SLS launch vehicles, as the vehicle, launch tower, and mobile launch platform were subjected to ground winds from all directions at varying heights for the vehicle off the launch pad. Additionally, aerodynamic forces and moments were acquired for all three SLS vehicles during the transition phase from liftoff to ascent that covered a wide range of angles of attack and angles of side slip. Details of the experimental setup including improved testing methods and a stiffer sting-balance system based on lessons learned from previous test entries are presented. Also, a new force measurement technique was applied during the test where two subminiature six-component load cells were installed in each Solid Rocket Booster (SRB) to acquire forces and moments for each SRB separately from the full vehicle forces and moments measured by the main strain gauge balance. Finally, sample results from the experiment are presented including improved overall data coverage, evaluation of the new SRB data, and smoke flow visualization photos.

Chan, David T.↗
NASA NTRS1993

Development of computational methods for heavy lift launch vehicles

The research effort has been focused on the development of an advanced flow solver for complex viscous turbulent flows with shock waves. The three-dimensional Euler and full/thin-layer Reynolds-averaged Navier-Stokes equations for compressible flows are solved on structured hexahedral grids. The Baldwin-Lomax algebraic turbulence model is used for closure. The space discretization is based on a cell-centered finite-volume method augmented by a variety of numerical dissipation models with optional total variation diminishing limiters. The governing equations are integrated in time by an implicit method based on lower-upper factorization and symmetric Gauss-Seidel relaxation. The algorithm is vectorized on diagonal planes of sweep using two-dimensional indices in three dimensions. A new computer program named CENS3D has been developed for viscous turbulent flows with discontinuities. Details of the code are described in Appendix A and Appendix B. With the developments of the numerical algorithm and dissipation model, the simulation of three-dimensional viscous compressible flows has become more efficient and accurate. The results of the research are expected to yield a direct impact on the design process of future liquid fueled launch systems.

Yoon, Seokkwan↗
NASA NTRS1965

A Method of Implementing Cutoff Conditions for Saturn V Lunar Missions Out of Earth Parking Orbit Assuming a Continuous Ground Launch Window

A method of implementing Saturn V lunar missions from an earth parking orbit is presented. The ground launch window is assumed continuous over a four and one-half hour period. The iterative guidance scheme combined with a set of auxiliary equations that define suitable S-IVB cutoff conditions, is the approach taken. The four inputs to the equations that define cutoff conditions are represented as simple third-degree polynomials as a function of ignition time. Errors at lunar arrival caused by the separate and combined effects of the guidance equations, cutoff conditions, hypersurface errors, and input representations are shown. Vehicle performance variations and parking orbit injection errors are included as perturbations. Appendix I explains how aim vectors were computed for the cutoff equations. Appendix II presents all guidance equations and related implementation procedures. Appendix III gives the derivation of the auxiliary cutoff equations. No error at lunar arrival was large enough to require a midcourse correction greater than one meter per second assuming a transfer time of three days and the midcourse correction occurs five hours after injection. Since this result is insignificant when compared to expected hardware errors, the implementation procedures presented are adequate to define cutoff conditions for Saturn V lunar missions.

PARKING ORBIT↗
NASA NTRS2001

Evaluation of Advanced Thermal Protection Techniques for Future Reusable Launch Vehicles

A method for integrating Aeroheating analysis into conceptual reusable launch vehicle RLV design is presented in this thesis. This process allows for faster turn-around time to converge a RLV design through the advent of designing an optimized thermal protection system (TPS). It consists of the coupling and automation of four computer software packages: MINIVER, TPSX, TCAT and ADS. MINIVER is an Aeroheating code that produces centerline radiation equilibrium temperatures, convective heating rates, and heat loads over simplified vehicle geometries. These include flat plates and swept cylinders that model wings and leading edges, respectively. TPSX is a NASA Ames material properties database that is available on the World Wide Web. The newly developed Thermal Calculation Analysis Tool (TCAT) uses finite difference methods to carry out a transient in-depth I-D conduction analysis over the center mold line of the vehicle. This is used along with the Automated Design Synthesis (ADS) code to correctly size the vehicle's thermal protection system JPS). The numerical optimizer ADS uses algorithms that solve constrained and unconstrained design problems. The resulting outputs for this process are TPS material types, unit thicknesses, and acreage percentages. TCAT was developed for several purposes. First, it provides a means to calculate the transient in-depth conduction seen by the surface of the TPS material that protects a vehicle during ascent and reentry. Along with the in-depth conduction, radiation from the surface of the material is calculated along with the temperatures at the backface and interior parts of the TPS material. Secondly, TCAT contributes added speed and automation to the overall design process. Another motivation in the development of TCAT is optimization.

Olds, John R.↗
NASA NTRS2003

Launch Vehicle Design and Optimization Methods and Priority for the Advanced Engineering Environment

NASA's Advanced Engineering Environment (AEE) is a research and development program that will improve collaboration among design engineers for launch vehicle conceptual design and provide the infrastructure (methods and framework) necessary to enable that environment. In this paper, three major technical challenges facing the AEE program are identified, and three specific design problems are selected to demonstrate how advanced methods can improve current design activities. References are made to studies that demonstrate these design problems and methods, and these studies will provide the detailed information and check cases to support incorporation of these methods into the AEE. This paper provides background and terminology for discussing the launch vehicle conceptual design problem so that the diverse AEE user community can participate in prioritizing the AEE development effort.

Rowell, Lawrence F.↗
NASA NTRSDate not supplied

Launch Vehicle Design and Requirements Verification Using Statistical Methods

- This presentation applies statistics to launch vehicle design, but the methods may be used for other engineering applications - Vehicle Models: What do we know when? How to separate parameters for making Monte Carlo runs - Known when assembling the vehicle but not during design - Known prior to committing to flight - Unknown at lift off - How do we correctly model the vehicle during the design phases? - Number of Monte Carlo samples - Requirements success - Design parameter values

Monte Carlo↗
NASA NTRS1992

Flowfield predictions for multiple body launch vehicles

A method is developed for simulating inviscid and viscous flow around multicomponent launch vehicles. Grids are generated by the GENIE general-purpose grid-generation code, and the flow solver is a finite-volume Runge-Kutta time-stepping method. Turbulence effects are simulated using Baldwin and Lomax (1978) turbulence model. Calculations are presented for three multibody launch vehicle configurations: one with two small-diameter solid motors, one with nine small-diameter solid motors, and one with three large-diameter solid motors.

Deese, Jerry E.↗
NASA NTRS2007

Aerial Videography From Locally Launched Rockets

A method of quickly collecting digital imagery of ground areas from video cameras carried aboard locally launched rockets has been developed. The method can be used, for example, to record rare or episodic events or to gather image data to guide decisions regarding treatment of agricultural fields or fighting wildfires. The method involves acquisition and digitization of a video frame at a known time along with information on the position and orientation of the rocket and camera at that time. The position and orientation data are obtained by use of a Global Positioning System receiver and a digital magnetic compass carried aboard the rocket. These data are radioed to a ground station, where they are processed, by a real-time algorithm, into georeferenced position and orientation data. The algorithm also generates a file of transformation parameters that account for the variation of image magnification and distortion associated with the position and orientation of the camera relative to the ground scene depicted in the image. As the altitude, horizontal position, and orientation of the rocket change between image frames, the algorithm calculates the corresponding new georeferenced position and orientation data and the associated transformation parameters. The output imagery can be rendered in any of a variety of formats. The figure presents an example of one such format.

Lyle, Stacey D.↗
NASA NTRS1983

Interplanetary mission design handbook. Volume 1, part 1: Earth to Venus ballistic mission opportunities, 1991-2005

Graphical data necessary for the preliminary design of ballistic missions to Venus is presented. Contours of launch energy requirements, as well as many other launch and arrival parameters, are presented in launch data/arrival date space for all launch opportunities from 1991 through 2005. An extensive text is included which explains mission design methods, from launch window development to Venus probe and orbiter arrival design, utilizing the graphical data in this volume as well as numerous equations relating various parameters.

Sergeyevsky, A. B.↗
NASA NTRS2010

Estimation of Supersonic Stage Separation Aerodynamics of Winged-Body Launch Vehicles Using Response Surface Methods

Response surface methodology was used to estimate the longitudinal stage separation aerodynamic characteristics of a generic, bimese, winged multi-stage launch vehicle configuration at supersonic speeds in the NASA LaRC Unitary Plan Wind Tunnel. The Mach 3 staging was dominated by shock wave interactions between the orbiter and booster vehicles throughout the relative spatial locations of interest. The inference space was partitioned into several contiguous regions within which the separation aerodynamics were presumed to be well-behaved and estimable using central composite designs capable of fitting full second-order response functions. The underlying aerodynamic response surfaces of the booster vehicle in belly-to-belly proximity to the orbiter vehicle were estimated using piecewise-continuous lower-order polynomial functions. The quality of fit and prediction capabilities of the empirical models were assessed in detail, and the issue of subspace boundary discontinuities was addressed. Augmenting the central composite designs to full third-order using computer-generated D-optimality criteria was evaluated. The usefulness of central composite designs, the subspace sizing, and the practicality of fitting lower-order response functions over a partitioned inference space dominated by highly nonlinear and possibly discontinuous shock-induced aerodynamics are discussed.

Erickson, Gary E.↗
NASA NTRS1989

A weak Hamiltonian finite element method for optimal guidance of an advanced launch vehicle

A temporal finite-element method based on a mixed form of the Hamiltonian weak principle is presented for optimal control problems. The mixed form of this principle contains both states and costates as primary variables, which are expanded in terms of nodal values and simple shape functions. Time derivatives of the states and costates do not appear in the governing variational equation; the only quantities whose time derivatives appear therein are virtual states and virtual costates. Numerical results are presented for an elementary trajectory optimization problem; they show very good agreement with the exact solution along with excellent computational efficiency and self-starting capability. The feasibility of this approach for real-time guidance applications is evaluated. A simplified model for an advanced launch vehicle application that is suitable for finite-element solution is presented.

Hodges, Dewey H.↗
NASA NTRS1995

Defining Support Requirements During Conceptual Design of Reusable Launch Vehicles

Current methods for defining the operational support requirements of new systems are data intensive and require significant design information. Methods are being developed to aid in the analysis process of defining support requirements for new launch vehicles during their conceptual design phase that work with the level of information available during this phase. These methods will provide support assessments based on the vehicle design and the operating scenarios. The results can be used both to define expected support requirements for new launch vehicle designs and to help evaluate the benefits of using new technologies. This paper describes the models, their current status, and provides examples of their use.

Morris, W. D.↗
NASA NTRS1982

Interplanetary mission design handbook. Volume 1, part 3: Earth to Jupiter ballistic mission opportunities, 1985-2005

Graphical data necessary for the preliminary design of ballistic missions to Jupiter are provided. Contours of launch energy requirements, as well as many other launch and Jupiter arrival parameters, are presented in launch date/arrival date space for all launch opportunities from 1985 through 2005. In addition, an extensive text is included which explains mission design methods, from launch window development to Jupiter probe and orbiter arrival design, utilizing the graphical data in this volume as well as numerous equations relating various parameters.

Sergeyevsky, A. B.↗
NASA NTRS1981

Interplanetary mission design handbook. Volume 1, part 4: Earth to Saturn ballistic mission opportunities, 1985-2005

Graphical data necessary for the preliminary design of ballistic missions to Saturn are provided. Contours of launch energy requirements as well as many other launch and Saturn arrival parameters, are presented in launch date/arrival date space for all launch opportunities from 1985 through 2005. In addition, an extensive text is included which explains mission design methods, from launch window development to Saturn probe and orbiter arrival design, utilizing the graphical data in this volume as well as numerous equations elating various parameters. This is the first of a planned series of mission design documents which will apply to all planets and some other bodies in the solar system.

Sergeyevsky, A. B.↗
NASA NTRS1983

Interplanetary mission design handbook. Volume 1, part 2: Earth to Mars ballistic mission opportunities, 1990-2005

Graphical data necessary for the preliminary design of ballistic missions to Mars are provided. Contours of launch energy requirements, as well as many other launch and Mars arrival parameters, are presented in launch date/arrival date space for all launch opportunities from 1990 through 2005. In addition, an extensive text is included which explains mission design methods, from launch window development to Mars probe and orbiter arrival design, utilizing the graphical data as well as numerous equations relating various parameters.

Sergeyevsky, A. B.↗