Booster attitude stabilization network synthesis Summary report
Saturn booster-attitude stabilization network synthesis - transfer function factoring methods and impedance synthesis techniques
TRANSFER FUNCTION↗
SEARCH · Search NASA
Results for “Launch Methods”
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.
Establishing Cleaning Methods for Cleanroom and Safety Suits at Remote Sites with Available Resources
"The James Webb Space Telescope (JWST) spent considerable time in multiple clean facilities during its launch campaign at the Centre Spatial Guyanais (CSG) in French Guiana. Throughout this time, it was imperative that personnel wear approved cleanroom or safety suits when working in proximity to the JWST observatory. This proved to be challenging for two reasons: 1) A large quantity of NASA cleanroom suits needed to be shipped to CSG to account for the heavy volume of NASA and ESA personnel requiring access to the clean facilities. It became evident that shipping clean suits from the US to CSG in a timely fashion during the launch campaign would be challenging. Consequently, a backup plan needed to be established to avoid running out. 2) For safety purposes, Self-Contained Atmospheric Protection Ensemble (SCAPE) and Splash suits were required for hazardous operations. Neither type of suit was cleanroom compatible; therefore, a viable cleaning process needed to be developed to ensure both types of suits met JWST’s approved cleanliness standards. The NASA Contamination Control (CC) team partnered with CSG’s S5E cleanroom and suit processing team. The teams were able to make modifications to the S5E facility to create a NASA approved makeshift clean environment where NASA cleanroom suits and CSG Splash suits were inspected, folded, and bagged for cleanroom use once they were washed and dried. Since SCAPE suits could not be cleaned using conventional methods, a procedure was developed to clean and inspect the suits by the CC team in a temporary clean environment prior to fueling. "
James Webb Space Telescope↗
Establishing Cleaning Methods For Cleanroom And Safety Suits At Remote Sites With Available Resources
The James Webb Space Telescope (JWST) spent considerable time in multiple clean facilities during its launch campaign at the Centre Spatial Guyanais (CSG) in French Guiana. Throughout this time, it was imperative that personnel wear approved cleanroom or safety suits when working in proximity to the JWST observatory. This proved to be challenging for two reasons: 1) A large quantity of NASA cleanroom suits needed to be shipped to CSG to account for the heavy volume of NASA and ESA personnel requiring access to the clean facilities. It became evident that shipping clean suits from the US to CSG in a timely fashion during the launch campaign would be challenging. Consequently, a backup plan needed to be established to avoid running out. 2) For safety purposes, Self-Contained Atmospheric Protection Ensemble (SCAPE) and Splash suits were required for hazardous operations. Neither type of suit was cleanroom compatible; therefore, a viable cleaning process needed to be developed to ensure both types of suits met JWST’s approved cleanliness standards. The NASA Contamination Control (CC) team partnered with CSG’s S5E cleanroom and suit processing team. The teams were able to make modifications to the S5E facility to create a NASA approved makeshift clean environment where NASA cleanroom suits and CSG Splash suits were inspected, folded, and bagged for cleanroom use once they were washed and dried. Since SCAPE suits could not be cleaned using conventional methods, a procedure was developed to clean and inspect the suits by the CC team in a temporary clean environment prior to fueling.
James Webb Space Telescope↗
Sensitivity Analysis of Launch Vehicle Debris Risk Model
As part of an analysis of the loss of crew risk associated with an ascent abort system for a manned launch vehicle, a model was developed to predict the impact risk of the debris resulting from an explosion of the launch vehicle on the crew module. The model consisted of a debris catalog describing the number, size and imparted velocity of each piece of debris, a method to compute the trajectories of the debris and a method to calculate the impact risk given the abort trajectory of the crew module. The model provided a point estimate of the strike probability as a function of the debris catalog, the time of abort and the delay time between the abort and destruction of the launch vehicle. A study was conducted to determine the sensitivity of the strike probability to the various model input parameters and to develop a response surface model for use in the sensitivity analysis of the overall ascent abort risk model. The results of the sensitivity analysis and the response surface model are presented in this paper.
Gee, Ken↗
Signal Processing Methods for Removing the Effects of Whole Body Vibration upon Speech
Humans may be exposed to whole-body vibration in environments where clear speech communications are crucial, particularly during the launch phases of space flight and in high-performance aircraft. Prior research has shown that high levels of vibration cause a decrease in speech intelligibility. However, the effects of whole-body vibration upon speech are not well understood, and no attempt has been made to restore speech distorted by whole-body vibration. In this paper, a model for speech under whole-body vibration is proposed and a method to remove its effect is described. The method described reduces the perceptual effects of vibration, yields higher ASR accuracy scores, and may significantly improve intelligibility. Possible applications include incorporation within communication systems to improve radio-communication systems in environments such a spaceflight, aviation, or off-road vehicle operations.
human response to vibration↗
Time-Domain Stability Margin Assessment
The baseline stability margins for NASA's Space Launch System (SLS) launch vehicle were generated via the classical approach of linearizing the system equations of motion and determining the gain and phase margins from the resulting frequency domain model. To improve the fidelity of the classical methods, the linear frequency domain approach can be extended by replacing static, memoryless nonlinearities with describing functions. This technique, however, does not address the time varying nature of the dynamics of a launch vehicle in flight. An alternative technique for the evaluation of the stability of the nonlinear launch vehicle dynamics along its trajectory is to incrementally adjust the gain and/or time delay in the time domain simulation until the system exhibits unstable behavior. This technique has the added benefit of providing a direct comparison between the time domain and frequency domain tools in support of simulation validation.
Clements, Keith↗
Time Domain Stability Margin Assessment of the NS Space Launch System GN&C Design for Exploration Mission One
The baseline stability margins for NASA's Space Launch System (SLS) launch vehicle were generated via the classical approach of linearizing the system equations of motion and determining the gain and phase margins from the resulting frequency domain model. To improve the fidelity of the classical methods, the linear frequency domain approach can be extended by replacing static, memoryless nonlinearities with describing functions. This technique, however, does not address the time varying nature of the dynamics of a launch vehicle in flight. An alternative technique for the evaluation of the stability of the nonlinear launch vehicle dynamics along its trajectory is to incrementally adjust the gain and/or time delay in the time domain simulation until the system exhibits unstable behavior. This technique has the added benefit of providing a direct comparison between the time domain and frequency domain tools in support of simulation validation.
Clements, Keith↗
Time Domain Stability Margin Assessment of the NASA Space Launch System GN&C Design for Exploration Mission One
The baseline stability margins for NASA's Space Launch System (SLS) launch vehicle were generated via the classical approach of linearizing the system equations of motion and determining the gain and phase margins from the resulting frequency domain model. To improve the fidelity of the classical methods, the linear frequency domain approach can be extended by replacing static, memoryless nonlinearities with describing functions. This technique, however, does not address the time varying nature of the dynamics of a launch vehicle in flight. An alternative technique for the evaluation of the stability of the nonlinear launch vehicle dynamics along its trajectory is to incrementally adjust the gain and/or time delay in the time domain simulation until the system exhibits unstable behavior. This technique has the added benefit of providing a direct comparison between the time domain and frequency domain tools in support of simulation validation.
Clements, Keith↗
Development of Buffet Forcing Functions Using Frequency-Dependent Coherence Factors
The current accepted approach to modeling launch vehicle transonic buffet environments is to acquire time-correlated unsteady pressure measurements at discrete locations on a model-scale wind-tunnel model and use these measurements to develop buffet forcing functions (BFFs). Part of the BFF development process is the application of coherence factors to account for the discrete nature of the pressure measurement used in the development of the BFFs. Presently, the Space Launch System (SLS) program divides the launch vehicle into distinct aerodynamic regions, within which, the coherence lengths are assumed to be constant. The coherence factors are computed by averaging the coherence function between sensors within the region over a specified frequency range. The present work validates and examines the impact of two proposed changes to the development of longitudinal coherence factors used in the development of launch vehicle BFFs. One change is to employ frequency-dependent coherence factors instead of coherence factors based on the mean of the coherence function. The second proposed change replaces the aerodynamic regions with a moving-segment approach that varies the calculated coherence lengths as a function of longitudinal location of the transducers. The impact of these approaches is examined using data from two rigid buffet model wind-tunnel tests: (1) a notional launch vehicle geometry through the comparison of discrete measurement-based BFFs to loads developed by continuous integration of unsteady pressure sensitive paint data and (2) the SLS Block 1 Cargo vehicle configuration for which BFFs have been previously developed using less-mature methods. The trends from this examination of updated coherence factor approaches ultimately result in more intuitive results than currently-accepted coherence methods.
buffet↗
Time-Accurate Computational Fluid Dynamics Simulation of a Pair of Moving Solid Rocket Boosters
Since the Columbia accident, the threat to the Shuttle launch vehicle from debris during the liftoff timeframe has been assessed by the Liftoff Debris Team at NASA/MSFC. In addition to engineering methods of analysis, CFD-generated flow fields during the liftoff timeframe have been used in conjunction with 3-DOF debris transport methods to predict the motion of liftoff debris. Early models made use of a quasi-steady flow field approximation with the vehicle positioned at a fixed location relative to the ground; however, a moving overset mesh capability has recently been developed for the Loci/CHEM CFD software which enables higher-fidelity simulation of the Shuttle transient plume startup and liftoff environment. The present work details the simulation of the launch pad and mobile launch platform (MLP) with truncated solid rocket boosters (SRBs) moving in a prescribed liftoff trajectory derived from Shuttle flight measurements. Using Loci/CHEM, time-accurate RANS and hybrid RANS/LES simulations were performed for the timeframe T0+0 to T0+3.5 seconds, which consists of SRB startup to a vehicle altitude of approximately 90 feet above the MLP. Analysis of the transient flowfield focuses on the evolution of the SRB plumes in the MLP plume holes and the flame trench, impingement on the flame deflector, and especially impingment on the MLP deck resulting in upward flow which is a transport mechanism for debris. The results show excellent qualitative agreement with the visual record from past Shuttle flights, and comparisons to pressure measurements in the flame trench and on the MLP provide confidence in these simulation capabilities.
Strutzenberg, Louise L.↗
A generalized modal shock spectra method for spacecraft loads analysis
Unlike the traditional shock spectra approach, the generalization presented in this paper permits elastic interaction between the spacecraft and launch vehicle in order to obtain accurate bounds on the spacecraft response and structural loads. In addition, the modal response from a previous launch vehicle transient analysis - with or without a dummy spacecraft - is exploited in order to define a modal impulse as a simple idealization of the actual forcing function. The idealized modal forcing function is then used to derive explicit expressions for an estimate of the bound on the spacecraft structural response and forces.
Trubert, M.↗
A method for simulating the atmospheric entry of long-range ballistic missiles
It is demonstrated with the aid of similitude arguments that a model launched from a hypervelocity gun upstream through a special supersonic nozzle should experience aerodynamic heating and resulting thermal stresses like those encountered by a long-range ballistic missile entering the earth's atmosphere. This demonstration hinges on the requirements that model and missile be geometrically similar and made of the same material, and that they have the same flight speed and Reynolds number (based on conditions just outside the boundary layer) at corresponding points in their trajectories. The hypervelocity gun provides the model with the required initial speed, while the nozzle scales the atmosphere, in terms of density variation, to provide the model with speeds and Reynolds numbers over its entire trajectory. Since both the motion and aerodynamic heating of a missile tend to be simulated in the model tests, this combination of hypervelocity gun and supersonic nozzle is termed an atmosphere entry simulator.
Eggers, A J , Jr↗
GTX Reference Vehicle Structural Verification Methods and Weight Summary
The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminary sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spread-sheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.
Hunter, J. E.↗
GTX Reference Vehicle Structural Verification Methods and Weight Summary
The design of a single-stage-to-orbit air breathing propulsion system requires the simultaneous development of a reference launch vehicle in order to achieve the optimal mission performance. Accordingly, for the GTX study a 300-lb payload reference vehicle was preliminarily sized to a gross liftoff weight (GLOW) of 238,000 lb. A finite element model of the integrated vehicle/propulsion system was subjected to the trajectory environment and subsequently optimized for structural efficiency. This study involved the development of aerodynamic loads mapped to finite element models of the integrated system in order to assess vehicle margins of safety. Commercially available analysis codes were used in the process along with some internally developed spreadsheets and FORTRAN codes specific to the GTX geometry for mapping of thermal and pressure loads. A mass fraction of 0.20 for the integrated system dry weight has been the driver for a vehicle design consisting of state-of-the-art composite materials in order to meet the rigid weight requirements. This paper summarizes the methodology used for preliminary analyses and presents the current status of the weight optimization for the structural components of the integrated system.
Hunter, J. E.↗
Space Launch System Complex Decision-Making Process
The Space Shuttle program has ended and elements of the Constellation Program have either been cancelled or transitioned to new NASA exploration endeavors. The National Aeronautics and Space Administration (NASA) has worked diligently to select an optimum configuration for the Space Launch System (SLS), a heavy lift vehicle that will provide the foundation for future beyond low earth orbit (LEO) large-scale missions for the next several decades. From Fall 2010 until Spring 2011, an SLS decision-making framework was formulated, tested, fully documented, and applied to multiple SLS vehicle concepts at NASA from previous exploration architecture studies. This was a multistep process that involved performing figure of merit (FOM)-based assessments, creating Pass/Fail gates based on draft threshold requirements, performing a margin-based assessment with supporting statistical analyses, and performing sensitivity analysis on each. This paper focuses on the various steps and methods of this process (rather than specific data) that allowed for competing concepts to be compared across a variety of launch vehicle metrics in support of the successful completion of the SLS Mission Concept Review (MCR) milestone.
Lyles, Garry↗
A Robust Method to Integrate End-to-End Mission Architecture Optimization Tools
End-to-end mission simulations include multiple phases of flight. For example, an end-to-end Mars mission simulation may include launch from Earth, interplanetary transit to Mars and entry, descent and landing. Each phase of flight is optimized to meet specified constraints and often depend on and impact subsequent phases. The design and optimization tools and methodologies used to combine different aspects of end-to-end framework and their impact on mission planning are presented. This work focuses on a robust implementation of a Multidisciplinary Design Analysis and Optimization (MDAO) method that offers the flexibility to quickly adapt to changing mission design requirements. Different simulations tailored to the liftoff, ascent, and atmospheric entry phases of a trajectory are integrated and optimized in the MDAO program Isight, which provides the user a graphical interface to link simulation inputs and outputs. This approach provides many advantages to mission planners, as it is easily adapted to different mission scenarios and can improve the understanding of the integrated system performance within a particular mission configuration. A Mars direct entry mission using the Space Launch System (SLS) is presented as a generic end-to-end case study. For the given launch period, the SLS launch performance is traded for improved orbit geometry alignment, resulting in an optimized a net payload that is comparable to that in the SLS Mission Planner's Guide.
Lugo, Rafael↗
Application of Kalman filtering techniques to strapdown system inertial alignment
Digital methods for determining initial alignment of strapdown inertial navigation system from vibration and sway corrupted data on launch pad
Earle B. Crocker↗
Effects of fragments
Methods used to determine structural damage caused by fragments from accidental explosions are outlined. Structures considered include launch facilities for liquid propellant rockets and metal plates and sheets. Penetrating and nonpenetrating fragments were considered.
Source record↗