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At least 649 records · Page 36

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NASA NTRS2001

Hyper-X Flight Engine Ground Testing for X-43 Flight Risk Reduction

Airframe-integrated scramjet engine testing has been completed at Mach 7 flight conditions in the NASA Langley 8-Foot High Temperature Tunnel as part of the NASA Hyper-X program. This test provided engine performance and operability data, as well as design and database verification, for the Mach 7 flight tests of the Hyper-X research vehicle (X-43), which will provide the first-ever airframe-integrated scramjet data in flight. The Hyper-X Flight Engine, a duplicate Mach 7 X-43 scramjet engine, was mounted on an airframe structure that duplicated the entire three-dimensional propulsion flowpath from the vehicle leading edge to the vehicle trailing edge. This model was also tested to verify and validate the complete flight-like engine system. This paper describes the subsystems that were subjected to flight-like conditions and presents supporting data. The results from this test help to reduce risk for the Mach 7 flights of the X-43.

Huebner, Lawrence D.↗

NASA NTRS2002

ISO 9000 and/or Systems Engineering Capability Maturity Model?

For businesses and organizations to remain competitive today they must have processes and systems in place that will allow them to first identify customer needs and then develop products/processes that will meet or exceed the customers needs and expectations. Customer needs, once identified, are normally stated as requirements. Designers can then develop products/processes that will meet these requirements. Several functions, such as quality management and systems engineering management are used to assist product development teams in the development process. Both functions exist in all organizations and both have a similar objective, which is to ensure that developed processes will meet customer requirements. Are efforts in these organizations being duplicated? Are both functions needed by organizations? What are the similarities and differences between the functions listed above? ISO 9000 is an international standard of goods and services. It sets broad requirements for the assurance of quality and for management's involvement. It requires organizations to document the processes and to follow these documented processes. ISO 9000 gives customers assurance that the suppliers have control of the process for product development. Systems engineering can broadly be defined as a discipline that seeks to ensure that all requirements for a system are satisfied throughout the life of the system by preserving their interrelationship. The key activities of systems engineering include requirements analysis, functional analysis/allocation, design synthesis and verification, and system analysis and control. The systems engineering process, when followed properly, will lead to higher quality products, lower cost products, and shorter development cycles. The System Engineering Capability Maturity Model (SE-CMM) will allow companies to measure their system engineering capability and continuously improve those capabilities. ISO 9000 and SE-CMM seem to have a similar objective, which is to document the organization's processes and certify to potential customers the capability of a supplier to control the processes that determine the quality of the product or services being produced. The remaining sections of this report examine the differences and similarities between ISO 9000 and SE-CMM and make recommendations for implementation.

Sampson E. Gholston↗

NASA NTRS2006

Certification Testing Approach for Propulsion System Design

The Certification of Propulsion Systems is costly and complex, involving development and qualification testing. The desire of the certification process is to assure all requirements can be demonstrated to be compliant. The purpose of this paper is to address the technical design concerns of certifying a propulsion system for flight. Presented are Pressurization, Tankage, Feed System and Combustion Instability concerns. Propulsion System Engineers are challenged with the dilemma for testing new systems to specific levels to reduce risk yet maintain budgetary targets. A methodical approach is presented to define the types of test suitable to address the technical issues for qualifying systems for retiring the risk levels. Experience of the lessons learned from supporting the Shuttle Program for Main Propulsion and On Orbit Propulsions Systems as well as previous collaborations on design concerns for certifying propulsion systems are utilized to address design concerns and verification approaches.

Rodriguez, Henry↗

NASA NTRS2006

Collected Papers in Structural Mechanics Honoring Dr. James H. Starnes, Jr.

This special publication contains a collection of structural mechanics papers honoring Dr. James H. Starnes, Jr. presented at the 46th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference held in Austin, Texas, April 18-21, 2005. Contributors to this publication represent a small number of those influenced by Dr. Starnes' technical leadership, his technical prowess and diversity, and his technical breath and depth in engineering mechanics. These papers cover some of the research areas Dr. Starnes investigated, which included buckling, postbuckling, and collapse of structures; composite structural mechanics, residual strength and damage tolerance of metallic and composite structures; and aircraft structural design, certification and verification. He actively pursued technical understanding and clarity, championed technical excellence, and modeled humility and perseverance.

Knight, Norman F., Jr.↗

NASA NTRS2008

Optimized Non-Obstructive Particle Damping (NOPD) Treatment for Composite Honeycomb Structures

Non-Obstructive Particle Damping (NOPD) technology is a passive vibration damping approach whereby metallic or non-metallic particles in spherical or irregular shapes, of heavy or light consistency, and even liquid particles are placed inside cavities or attached to structures by an appropriate means at strategic locations, to absorb vibration energy. The objective of the work described herein is the development of a design optimization procedure and discussion of test results for such a NOPD treatment on honeycomb (HC) composite structures, based on finite element modeling (FEM) analyses, optimization and tests. Modeling and predictions were performed and tests were carried out to correlate the test data with the FEM. The optimization procedure consisted of defining a global objective function, using finite difference methods, to determine the optimal values of the design variables through quadratic linear programming. The optimization process was carried out by targeting the highest dynamic displacements of several vibration modes of the structure and finding an optimal treatment configuration that will minimize them. An optimal design was thus derived and laboratory tests were conducted to evaluate its performance under different vibration environments. Three honeycomb composite beams, with Nomex core and aluminum face sheets, empty (untreated), uniformly treated with NOPD, and optimally treated with NOPD, according to the analytically predicted optimal design configuration, were tested in the laboratory. It is shown that the beam with optimal treatment has the lowest response amplitude. Described below are results of modal vibration tests and FEM analyses from predictions of the modal characteristics of honeycomb beams under zero, 50% uniform treatment and an optimal NOPD treatment design configuration and verification with test data.

Panossian, H.↗

NASA NTRS2008

Surface Landing Site Weather Analysis for NASA's Constellation Program

Altino, Karen M.↗

NASA NTRS2008

Database Driven 6-DOF Trajectory Simulation for Debris Transport Analysis

Debris mitigation and risk assessment have been carried out by NASA and its contractors supporting Space Shuttle Return-To-Flight (RTF). As a part of this assessment, analysis of transport potential for debris that may be liberated from the vehicle or from pad facilities prior to tower clear (Lift-Off Debris) is being performed by MSFC. This class of debris includes plume driven and wind driven sources for which lift as well as drag are critical for the determination of the debris trajectory. As a result, NASA MSFC has a need for a debris transport or trajectory simulation that supports the computation of lift effect in addition to drag without the computational expense of fully coupled CFD with 6-DOF. A database driven 6-DOF simulation that uses aerodynamic force and moment coefficients for the debris shape that are interpolated from a database has been developed to meet this need. The design, implementation, and verification of the database driven six degree of freedom (6-DOF) simulation addition to the Lift-Off Debris Transport Analysis (LODTA) software are discussed in this paper.

West, Jeff↗

NASA NTRS2008

Surface Landing Site Weather Analysis for Constellation Program

Weather information is an important asset for NASA's Constellation Program in developing the next generation space transportation system to fly to the International Space Station, the Moon and, eventually, to Mars. Weather conditions can affect vehicle safety and performance during multiple mission phases ranging from pre-launch ground processing to landing and recovery operations, including all potential abort scenarios. Meteorological analysis is an important contributor, not only to the development and verification of system design requirements but also to mission planning and active ground operations. Of particular interest are the surface atmospheric conditions at both nominal and abort landing sites for the manned Orion capsule. Weather parameters such as wind, rain, and fog all play critical roles in the safe landing of the vehicle and subsequent crew and vehicle recovery. The Marshall Space Flight Center Natural Environments Branch has been tasked by the Constellation Program with defining the natural environments at potential landing zones. Climatological time series of operational surface weather observations are used to calculate probabilities of occurrence of various sets of hypothetical vehicle constraint thresholds, Data are available for numerous geographical locations such that statistical analysis can be performed for single sites as well as multiple-site network configurations. Results provide statistical descriptions of how often certain weather conditions are observed at the site(s) and the percentage that specified criteria thresholds are matched or exceeded. Outputs are tabulated by month and hour of day to show both seasonal and diurnal variation. This paper will describe the methodology used for data collection and quality control, detail the types of analyses performed, and provide a sample of the results that can be obtained,

Altino, Karen M.↗

NASA NTRS2008

NASA Experience with Pogo in Human Spaceflight Vehicles

An overview of more than 45 years of NASA human spaceflight experience is presented with respect to the thrust axis vibration response of liquid fueled rockets known as pogo. A coupled structure and propulsion system instability, pogo can result in the impairment of the astronaut crew, an unplanned engine shutdown, loss of mission, or structural failure. The NASA history begins with the Gemini Program and adaptation of the USAF Titan II ballistic missile as a spacecraft launch vehicle. It continues with the pogo experienced on several Apollo-Saturn flights in both the first and second stages of flight. The defining moment for NASA s subsequent treatment of pogo occurred with the near failure of the second stage on the ascent of the Apollo 13 mission. Since that time NASA has had a strict "no pogo" philosophy that was applied to the development of the Space Shuttle. The "no pogo" philosophy lead to the first vehicle designed to be pogo-free from the beginning and the first development of an engine with an integral pogo suppression system. Now, more than 30 years later, NASA is developing two new launch vehicles, the Ares I crew launch vehicle propelling the Orion crew excursion vehicle, and the Ares V cargo launch vehicle. A new generation of engineers must again exercise NASA s system engineering method for pogo mitigation during design, development and verification.

Larsen, Curtis E.↗

NASA NTRS2007

Programmable Thermostats for MPLM Shell Heater Control ULF1: Thermal Performances - 1

The Multi-Purpose Logistics Module (MPLM) is the primary carrier for "pressurized" logistics to and from the International Space Station (ISS). The MPLM is transported in the payload bay of the Space Shuttle and is docked to the ISS for unloading, and reloading, of contents within the ISS shirt sleeve environment. Foil heaters, controlled originally with bi-metallic thermostats, are distributed across the outside of the MPLM structure and are utilized to provide energy to the structure to avoid exposure to cold temperatures and prevent condensation. The existing bi-metallic, fixed temperature set point thermostats have been replaced with Programmable Thermostats Modules (PTMs) in the Passive Thermal Control Subsystem (PTCS) 28Vdc shell heater circuits. The goal of using the PTM thermostat is to improve operational efficiency of the MPLM on-orbit shell heaters by providing better shell temperature control via feedback control capability. Each heater circuit contains a programmable thermostat connected to an external temperature sensor, a Resistive Temperature Device (RTD), which is used to provide continuous temperature monitoring capability. Each thermostat has programmable temperature set points and control spans. The data acquisition system uses a standard RS-485 serial interface communications cable to provide digital control capability. The PTM system was designed by MSFC, relying upon ALTEC support for their integration within the MPLM system design, while KSC performed the installation and ground checkout testing of the thermostat and RS-485 communication cable on the MPLM FM1 flight module. The PTMs were used for the first time during the STS-121/ULF1.1 mission. This paper will describe the design, development and verification of the PTM system, as well as the PTM flight performance and comparisons with SINDA thermal model predictions.

Glasgow, Shaun↗

NASA NTRS2009

Engine Gimbal Requirements for Ground Testing of J-2X

Based on the Apollo-era J-2 that powered the second and third stages of the Saturn V, the current J-2X is the liquid hydrogen and oxygen high-altitude rocket engine in development for both the Ares I Upper Stage and Ares V Earth Departure Stage. During my summer 2009 internship, J-2X was at a stage in its design maturity where verification testing needed to be considered for the benefit of adequate test facility preparation. My task was to focus on gimbal requirements and gimbal related hot-fire test plans. Facility capabilities were also of interest, specifically for hot-fire testing slated to occur at test stands A-1, A-2, and A-3 at Stennis Space Center(SSC) in Bay St. Louis, Mississippi. Gimbal requirements and stage interface conditions were investigated by applying a top-to-bottom systems engineering approach, which involved system level requirements, engine level requirements from both government and engine contractor perspectives, component level requirements, and the J-2X to Upper Stage and Earth Departure Stage interface control documents. Previous hydrogen and oxygen liquid rocket engine gimbal verification methods were researched for a glimpse at lessons learned. Discussion among the J-2X community affected by gimballing was organized to obtain input relative to proper verification of their respective component. Implementing suggestions such as gimbal pattern, angulated dwell time, altitude testing options, power level, and feed line orientation, I was able to match tests to test stands in the A Complex at SSC. Potential test capability gaps and risks were identified and pursued. The culmination of all these efforts was to coordinate with SSC to define additional facility requirements for both the A-3 altitude test stand that is currently under construction and the A-1 sea level test stand which is being renovated

Kovalcik, Julia↗

NASA NTRS2009

Ares I-X Roll Control System Development

Project Managers often face challenging technical, schedule and budget issues. This presentation will explore how the Ares I-X Roll Control System Integrated Product Team (IPT) mitigated challenges such as concurrent engineering requirements and environments and evolving program processes, while successfully managing an aggressive project schedule and tight budget. IPT challenges also included communications and negotiations among inter- and intra-government agencies, including the US Air Force, NASA/MSFC Propulsion Engineering, LaRC, GRC, KSC, WSTF, and the Constellation Program. In order to successfully meet these challenges it was essential that the IPT define those items that most affected the schedule critical path, define early mitigation strategies to reduce technical, schedule, and budget risks, and maintain the end-product focus of an "unmanned test flight" context for the flight hardware. The makeup of the IPT and how it would function were also important considerations. The IPT consisted of NASA/MSFC (project management, engineering, and safety/quality) and contractors (Teledyne Brown Engineering and Pratt and Whitney Rocketdyne, who supplied heritage hardware experience). The early decision to have a small focused IPT working "badgelessly" across functional lines to eliminate functional stove-piping allowed for many more tasks to be done by fewer people. It also enhanced a sense of ownership of the products, while still being able to revert back to traditional roles in order to provide the required technical independence in design reviews and verification closures. This presentation will highlight several prominent issues and discuss how they were mitigated and the resulting Lessons Learned that might benefit other projects.

Unger, Ronald J.↗

NASA NTRS2009

Explicit Finite Element Modeling of Multilayer Composite Fabric for Gas Turbine Engine Containment Systems, Phase II: Material Model Development and Simulation of Experiments - Part 3

A team consisting of Arizona State University, Honeywell Engines, Systems & Services, the National Aeronautics and Space Administration Glenn Research Center, and SRI International collaborated to develop computational models and verification testing for designing and evaluating turbine engine fan blade fabric containment structures. This research was conducted under the Federal Aviation Administration Airworthiness Assurance Center of Excellence and was sponsored by the Aircraft Catastrophic Failure Prevention Program. The research was directed toward improving the modeling of a turbine engine fabric containment structure for an engine blade-out containment demonstration test required for certification of aircraft engines. The research conducted in Phase II began a new level of capability to design and develop fan blade containment systems for turbine engines. Significant progress was made in three areas: (1) further development of the ballistic fabric model to increase confidence and robustness in the material models for the Kevlar(TradeName) and Zylon(TradeName) material models developed in Phase I, (2) the capability was improved for finite element modeling of multiple layers of fabric using multiple layers of shell elements, and (3) large-scale simulations were performed. This report concentrates on the material model development and simulations of the impact tests.

Zylon↗

NASA NTRS2011

Comparison of Adhesion and Retention Forces for Two Candidate Docking Seal Elastomers

To successfully mate two pressurized vehicles or structures in space, advanced seals are required at the interface to prevent the loss of breathable air to the vacuum of space. A critical part of the development testing of candidate seal designs was a verification of the integrity of the retaining mechanism that holds the silicone seal component to the structure. Failure to retain the elastomer seal during flight could liberate seal material in the event of high adhesive loads during undocking. This work presents an investigation of the force required to separate the elastomer from its metal counter-face surface during simulated undocking as well as a comparison to that force which was necessary to destructively remove the elastomer from its retaining device. Two silicone elastomers, Wacker 007-49524 and Esterline ELASA-401, were evaluated. During the course of the investigation, modifications were made to the retaining devices to determine if the modifications improved the force needed to destructively remove the seal. The tests were completed at the expected operating temperatures of -50, +23, and +75 C. Under the conditions investigated, the comparison indicated that the adhesion between the elastomer and the metal counter-face was significantly less than the force needed to forcibly remove the elastomer seal from its retainer, and no failure would be expected.

Hartzler, Brad D.↗

NASA NTRS2011

Structural Design Requirements and Factors of Safety for Spaceflight Hardware: For Human Spaceflight

This document establishes the structural requirements for human-rated spaceflight hardware including launch vehicles, spacecraft and payloads. These requirements are applicable to Government Furnished Equipment activities as well as all related contractor, subcontractor and commercial efforts. These requirements are not imposed on systems other than human-rated spacecraft, such as ground test articles, but may be tailored for use in specific cases where it is prudent to do so such as for personnel safety or when assets are at risk. The requirements in this document are focused on design rather than verification. Implementation of the requirements is expected to be described in a Structural Verification Plan (SVP), which should describe the verification of each structural item for the applicable requirements. The SVP may also document unique verifications that meet or exceed these requirements with NASA Technical Authority approval.

Bernstein, Karen S.↗

NASA NTRS2012

Integrated Data Modeling and Simulation on the Joint Polar Satellite System Program

The Joint Polar Satellite System is a modern, large-scale, complex, multi-mission aerospace program, and presents a variety of design, testing and operational challenges due to: (1) System Scope: multi-mission coordination, role, responsibility and accountability challenges stemming from porous/ill-defined system and organizational boundaries (including foreign policy interactions) (2) Degree of Concurrency: design, implementation, integration, verification and operation occurring simultaneously, at multiple scales in the system hierarchy (3) Multi-Decadal Lifecycle: technical obsolesce, reliability and sustainment concerns, including those related to organizational and industrial base. Additionally, these systems tend to become embedded in the broader societal infrastructure, resulting in new system stakeholders with perhaps different preferences (4) Barriers to Effective Communications: process and cultural issues that emerge due to geographic dispersion and as one spans boundaries including gov./contractor, NASA/Other USG, and international relationships.

Roberts, Christopher J.↗

NASA NTRS2011

Fluid Structure Interaction of Parachutes in Supersonic Planetary Entry

A research program to provide physical insight into disk-gap-band parachute operation in the supersonic regime on Mars was conducted. The program included supersonic wind tunnel tests, computational fluid dynamics and fluid structure interaction simulations. Specifically, the nature and cause of the "area oscillation" phenomenon were investigated to determine the scale, aerodynamic, and aero-elastic dependence of the supersonic parachute collapse and re-inflation event. A variety of non-intrusive, temporally resolved, and high resolution diagnostic techniques were used to interrogate the flow and generate validation datasets. The results of flow visualization, particle image velocimetry, load measurements, and photogrammetric reconstruction will be presented. Implications to parachute design, use, and verification will also be discussed.

parachute↗

NASA NTRS2012

A Probabilistic Mass Estimation Algorithm for a Novel 7- Channel Capacitive Sample Verification Sensor

A document describes an algorithm created to estimate the mass placed on a sample verification sensor (SVS) designed for lunar or planetary robotic sample return missions. A novel SVS measures the capacitance between a rigid bottom plate and an elastic top membrane in seven locations. As additional sample material (soil and/or small rocks) is placed on the top membrane, the deformation of the membrane increases the capacitance. The mass estimation algorithm addresses both the calibration of each SVS channel, and also addresses how to combine the capacitances read from each of the seven channels into a single mass estimate. The probabilistic approach combines the channels according to the variance observed during the training phase, and provides not only the mass estimate, but also a value for the certainty of the estimate. SVS capacitance data is collected for known masses under a wide variety of possible loading scenarios, though in all cases, the distribution of sample within the canister is expected to be approximately uniform. A capacitance-vs-mass curve is fitted to this data, and is subsequently used to determine the mass estimate for the single channel s capacitance reading during the measurement phase. This results in seven different mass estimates, one for each SVS channel. Moreover, the variance of the calibration data is used to place a Gaussian probability distribution function (pdf) around this mass estimate. To blend these seven estimates, the seven pdfs are combined into a single Gaussian distribution function, providing the final mean and variance of the estimate. This blending technique essentially takes the final estimate as an average of the estimates of the seven channels, weighted by the inverse of the channel s variance.

Wolf, Michael↗