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At least 1,135 records · Page 63
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Launch Vehicle Control Center Architectures

This analysis is a survey of control center architectures of the NASA Space Launch System (SLS), United Launch Alliance (ULA) Atlas V and Delta IV, and the European Space Agency (ESA) Ariane 5. Each of these control center architectures have similarities in basic structure, and differences in functional distribution of responsibilities for the phases of operations: (a) Launch vehicles in the international community vary greatly in configuration and process; (b) Each launch site has a unique processing flow based on the specific configurations; (c) Launch and flight operations are managed through a set of control centers associated with each launch site, however the flight operations may be a different control center than the launch center; and (d) The engineering support centers are primarily located at the design center with a small engineering support team at the launch site.

Watson, Michael D.↗
NASA NTRS2006

Spitzer Pre Launch Mission Operations System - The Road to Launch

Spitzer Space Telescope was launched on 25 August 2003 into an Earth-trailing solar orbit to acquire infrared observations from space. Development of the Mission Operations System (MOS) portion prior to launch was very different from planetary missions from the stand point that the MOS teams and Ground Data System had to be ready to support all aspects of the mission at launch (i.e., no cruise period for finalizing the implementation). For Spitzer, all mission-critical events post launch happen in hours or days rather than months or years, as is traditional with deep space missions. At the end of 2000 the Project was dealt a major blow when the Mission Operations System (MOS) had an unsuccessful Critical Design Review (CDR). The project made major changes at the beginning of 2001 in an effort to get the MOS (and Project) back on track. The result for the Spitzer Space Telescope was a successful launch of the observatory followed by an extremely successful In Orbit Checkout (IOC) and operations phase. This paper describes how the project was able to recover the MOS to a successful Delta (CDR) by mid 2001, and what changes in philosophies, experiences, and lessons learned followed. It describes how projects must invest early or else invest heavily later in the development phase to achieve a successful operations phase.

observatory↗
NASA NTRS2015

Launch Vehicle Manual Steering with Adaptive Augmenting Control:In-Flight Evaluations of Adverse Interactions Using a Piloted Aircraft

An Adaptive Augmenting Control (AAC) algorithm for the Space Launch System (SLS) has been developed at the Marshall Space Flight Center (MSFC) as part of the launch vehicle's baseline flight control system. A prototype version of the SLS flight control software was hosted on a piloted aircraft at the Armstrong Flight Research Center to demonstrate the adaptive controller on a full-scale realistic application in a relevant flight environment. Concerns regarding adverse interactions between the adaptive controller and a potential manual steering mode were also investigated by giving the pilot trajectory deviation cues and pitch rate command authority, which is the subject of this paper. Two NASA research pilots flew a total of 25 constant pitch rate trajectories using a prototype manual steering mode with and without adaptive control, evaluating six different nominal and off-nominal test case scenarios. Pilot comments and PIO ratings were given following each trajectory and correlated with aircraft state data and internal controller signals post-flight.

adaptive control↗
NASA NTRS2014

Conceptual Launch Vehicle and Spacecraft Design for Risk Assessment

One of the most challenging aspects of developing human space launch and exploration systems is minimizing and mitigating the many potential risk factors to ensure the safest possible design while also meeting the required cost, weight, and performance criteria. In order to accomplish this, effective risk analyses and trade studies are needed to identify key risk drivers, dependencies, and sensitivities as the design evolves. The Engineering Risk Assessment (ERA) team at NASA Ames Research Center (ARC) develops advanced risk analysis approaches, models, and tools to provide such meaningful risk and reliability data throughout vehicle development. The goal of the project presented in this memorandum is to design a generic launch 7 vehicle and spacecraft architecture that can be used to develop and demonstrate these new risk analysis techniques without relying on other proprietary or sensitive vehicle designs. To accomplish this, initial spacecraft and launch vehicle (LV) designs were established using historical sizing relationships for a mission delivering four crewmembers and equipment to the International Space Station (ISS). Mass-estimating relationships (MERs) were used to size the crew capsule and launch vehicle, and a combination of optimization techniques and iterative design processes were employed to determine a possible two-stage-to-orbit (TSTO) launch trajectory into a 350-kilometer orbit. Primary subsystems were also designed for the crewed capsule architecture, based on a 24-hour on-orbit mission with a 7-day contingency. Safety analysis was also performed to identify major risks to crew survivability and assess the system's overall reliability. These procedures and analyses validate that the architecture's basic design and performance are reasonable to be used for risk trade studies. While the vehicle designs presented are not intended to represent a viable architecture, they will provide a valuable initial platform for developing and demonstrating innovative risk assessment capabilities.

Launch Vehicle↗
NASA NTRS2000

Commercial Space Policy in the 1980s: Proceedings of a Roundtable Discussion

The Space Business Archives and the NASA History Office signed a Memorandum of Understanding in March of 1999. The MOU outlines several opportunities for cooperative endeavors between the two agencies in historical programming. This oral history, and subsequently this publication, are the first products of that cooperation. In accordance with the purpose of the Space Business Archives--to provide an impartial forum for lessons learned in the development of the commercial space industry--the idea for this roundtable discussion seemed appropriate as the Archives first public program. With the combined resources of the Archives and the NASA History Office we were fortunate to assemble a panel of individuals that served in both industry and government during the 1980s, many working in both sectors during that time. When envisioning the focus of this oral history, we decided that it was appropriate to highlight space policy in the 1980s, with an emphasis on the emerging commercial industry. Panelists were sent several documents in preparation, such as the Land Remote Sensing Commercialization Act and the Commercial Space Launch Act of 1984, President Reagan's 1982 National Space Policy, and other memoranda and letters that outline important policy issues of the decade. This discussion, we think, fills in some of the gaps that would otherwise be left unfilled when simply reading through the documents themselves. Some of these gaps include: how were these policy directives, legislation and decisions introduced and developed, by whom, and at what political and financial cost? This transcript is meant to serve as a reference to some of the issues, organizations and individuals involved in the creation and development of space policy during the 1980s. It is also the result of the first of many future roundtable discussions aimed at providing an open exchange of ideas concerning past success and failure in order to provide a stronger base for future endeavors in governmental, civil and commercial cooperation in space.

Dahlstrom, Neil↗
NASA NTRS2004

2003 NASA Seal/Secondary Air System Workshop

The following reports were included in the 2003 NASA Seal/Secondary Air System Workshop:Low Emissions Alternative Power (LEAP); Overview of NASA Glenn Seal Developments; NASA Ultra Efficient Engine Technology Project Overview; Development of Higher Temperature Abradable Seals for Industrial Gas Turbines; High Misalignment Carbon Seals for the Fan Drive Gear System Technologies; Compliant Foil Seal Investigations; Test Rig for Evaluating Active Turbine Blade Tip Clearance Control Concepts; Controls Considerations for Turbine Active Clearance Control; Non-Contacting Finger Seal Developments and Design Considerations; Effect of Flow-Induced Radial Load on Brush Seal/Rotor Contact Mechanics; Seal Developments at Flowserve Corporation; Investigations of High Pressure Acoustic Waves in Resonators With Seal-Like Features; Numerical Investigations of High Pressure Acoustic Waves in Resonators; Feltmetal Seal Material Through-Flow; "Bimodal" Nuclear Thermal Rocket (BNTR) Propulsion for Future Human Mars Exploration Missions; High Temperature Propulsion System Structural Seals for Future Space Launch Vehicles; Advanced Control Surface Seal Development for Future Space Vehicles; High Temperature Metallic Seal Development for Aero Propulsion and Gas Turbine Applications; and BrazeFoil Honeycomb.

Steinetz, Bruce M.↗
NASA NTRS2005

Future Interagency Range and Spaceport Technologies (FIRST) Formulation Products: 1. Transformational Spaceport and Range Concept of Operations. 2. F.I.R.S.T. Business Case Analysis

The Baseline Report captures range and spaceport capabilities at five sites: KSC, CCAFS, VAFB, Wallops, and Kodiak. The Baseline depicts a future state that relies on existing technology, planned upgrades, and straight-line recapitalization at these sites projected through 2030. The report presents an inventory of current spaceport and range capabilities at these five sites. The baseline is the first part of analyzing a business case for a set of capabilities designed to transform U.S. ground and space launch operations toward a single, integrated national "system" of space transportation systems. The second part of the business case compares current capabilities with technologies needed to support the integrated national "system". The final part, a return on investment analysis, identifies the technologies that best lead to the integrated national system and reduce recurring costs..Numerous data sources were used to define and describe the baseline spaceport and range by identifying major systems and elements and describing capabilities, limitations, and capabilities

Source record↗
NASA NTRS2015

Attitude Ground System (AGS) for the Magnetospheric Multi-Scale (MMS) Mission

MMS Overview Recall from Conrads presentation earlier today MMS launch: March 13, 2015 on an Atlas V from Space Launch Complex 40, Cape Canaveral, Florida MMS Observatory Separation: five minute intervals spinning at 3 rpm approximately 1.5 hours after launch MMS Science Goals: study magnetospheric plasma physics and understand the processes that cause power grids, communication disruptions and Aurora formation Mission: 4 identical spacecraft in tetrahedral formation with variable size1.2 x 12 RE in Phase 1, with apogee on dayside to observe bow shock1.2 x 25 RE in Phase 2, with apogee on night side to observe magneto tail Challenges Tight attitude control box, orbit and formation maintenance requirements Maneuvers on thrusters every two weeks Delta-H Spin axis direction and spin rate maintenance Delta-V Orbit and Formation maintenance Mission phase transitions AGS support Smart targeting prediction of Spin-Axis attitude in the presence of environmental torques to stay within the science attitude Determination of the spacecraft attitude and spin rate (sensitive to knowledge of inertia tensor)Calibrations to improve attitude determination results and improve orbit maneuvers Mass properties (Center of Mass, and inertia tensor for nutation and coning) Accelerometer bias (sensitive to the accuracy of the rate estimates) Sensor alignments.

Estimation and Calibration Support↗
NASA NTRS2018

Analysis of Near Rectilinear Halo Orbit Insertion with a 40-kW Solar Electric Propulsion System

This paper examines low thrust trajectories for delivery of a 40-kW solar electric propulsion spacecraft and potential additional payload to a desired NRHO. One option considered is a trans-lunar injection launch as a co-manifested payload on the Space Launch System. For this option, a reference trajectory is designed and a scan of launch dates is completed to understand the propellant mass sensitivity. A 15-day period cyclical variation in required propellant is observed that is attributed to solar gravity effects. A second option considered is to launch on a smaller commercial launch vehicle to a less energetic elliptical orbit and use SEP to spiral out to NRHO. For this option, analysis is completed to understand the trades between delivered mass to NRHO, total propellant required, time of flight, and solar array degradation. Results show that, while launching to lower altitudes can deliver greater payload mass to NRHO, significant solar array degradation can be observed.

Trajectory↗
NASA NTRS2018

Analysis of Near Rectilinear Halo Orbit Insertion with a 40-kW Solar Electric Propulsion System

This paper examines low thrust trajectories for delivery of a 40-kW solar electric propulsion spacecraft and potential additional payload to a desired NRHO. One option considered is a trans-lunar injection launch as a co-manifested payload on the Space Launch System. For this option, a reference trajectory is designed and a scan of launch dates is completed to understand the propellant mass sensitivity. A 15-day period cyclical variation in required propellant is observed that is attributed to solar gravity effects. A second option considered is to launch on a smaller commercial launch vehicle to a less energetic elliptical orbit and use SEP to spiral out to NRHO. For this option, analysis is completed to understand the trades between delivered mass to NRHO, total propellant required, time of flight, and solar array degradation. Results show that, while launching to lower altitudes can deliver greater payload mass to NRHO, significant solar array degradation can be observed.

Mccarty, Steven L.↗
NASA NTRS2017

Telecommunication System Design for Interplanetary CubeSat Missions: LunaH-Map

The Lunar Hydrogen Mapper (Luna H-Map) will be one of 13 CubeSats to launch on the first integrated flight of NASA’s Space Launch System and Orion spacecraft in 2018. The goal of the LunaH-Map mission is to map the hydrogen content of the entire South Pole of the moon, including permanently shadowed regions at high resolution. The spacecraft is a 6 U CubeSat and the main instrument is a neutron spectrometer which will be used to perform the measurements required to accomplish the primary science objective. Once LunaH-Map reaches the Moon, the spacecraft will perform a 60 - day science mission, consisting of 141 science orbits during which the measurements will be taken. In the design of the LunaH-Map spacecraft, one of the main challenges is the elecommunication system as the spacecraft needs the ability to relay data during the long cruise phase and during the science phase. In addition, the telecommunication system needs to support the main navigation functions required to reach the Moon and to insert in the highly elliptical orbit required for the science phase. This paper covers the main aspects of the telecommunication design for the mission including: link and coverage analysis, waveform selection and spectral constraints, hardware selection and ground station coordination.

Babuscia, Alessandra↗
NASA NTRSDate not supplied

Lightning Protection System

A new Lightning Protection System (LPS) was designed and built at Launch Complex 39B (LC-39B), at the Kennedy Space Center (KSC), Florida, in 2009, for NASA’s Space Launch System (SLS) program. A comprehensive lightning monitoring system was additionally designed and the LC-39B lightning instrumentation started incremental deployment early 2011. By March 2011, the LC-39B lightning monitoring system was activated with short outages during construction and renovations of the LC-39B. Since 2011, the LC-39B LPS has been directly struck at least once per year. The earliest and latest in a year that the LPS has been directly struck by lightning has been January 7th and October 1st (both 2017), respectively. During the summer of 2018, the Mobile Launcher 1 (ML-1) lightning monitoring system was deployed and partially active during ML-1 construction/testing. Since 2018, the ML-1 has been parked at the LC-39B for different durations of time for integrated testing. Early August 2019 (6th and 9th ) multiple lightning strikes terminated directly on the LPS at three different locations in addition to a nearby strike within the Pad perimeter and multiple nearby strikes outside the Pad Perimeter, triggering both, the LC-39B and the ML-1 lightning monitoring systems. This has been the first time direct lightning strikes have terminated on the LPS or within the LC-39B perimeter while the ML-1 has been at the LC-39B. This paper summarizes direct and close nearby lightning strikes to LC-39B from 2011 until the summer of 2019.

Angel G Mata↗
NASA NTRS2022

Accelerance Decoupling: An Approach for Removing the Influence of the Test Stand from the Integrated Modal Test

The main objective for launch vehicle (LV) modal testing is to quantify the LV’s modal properties in the free-free state (post pad separation). However, given the size of most LV systems, free-free testing is a challenge and often not feasible. With this, a test stand, typically the launch pad itself, is introduced as the means of support. This shifts the challenge to developing robust numerical methods for removing the influence of the launch pad from the integrated system modal test. The Space Launch System (SLS) is no exception where the mobile launcher (ML) is used to support the vehicle for the integrated modal test (IMT). For the IMT, it is well understood from pre-test analysis with finite element models (FEMs) of the SLS and SLS coupled to ML that the ML has a significant influence on the SLS modal properties especially in the lower frequency range where the primary SLS bending modes exist. An accelerance decoupling (AD) method has been formulated for the purpose of “subtracting out” the influence of the ML from the IMT results. With AD, the SLS decoupled frequency response functions (FRFs) are directly extracted from the IMT FRFs. The subject approach is aimed to utilize measured data only and achieve a robust FRF decoupling scheme. AD is derived from a widely used coupling technique called “Receptance Coupling” (RC). The AD core equation reverses the RC process and utilizes a pair of auxiliary equations that enable the core equation to be resolved based on measured data only. In AD, the decoupled component FRFs are extracted from the coupled system FRFs with a transformation to remove the contribution of the “subtractive component”. This paper addresses the AD’s operational flexibility to resolve SLS free-free modal properties from coupled system measured data but also the possibility to include data from FEM if there is enough confidence in the FEM or if it is asserted that the effect to the final outcome is reasonable.

Accelerance decoupling↗
NASA NTRS2022

Accelerance Decoupling: An Approach for Removing the Influence of the Test Stand from the Integrated Modal Test

The main objective for launch vehicle (LV) modal testing is to quantify the LV’s modal properties in the free-free state (post pad separation). However, given the size of most LV systems, free-free testing is a challenge and often not feasible. With this, a test stand, typically the launch pad itself, is introduced as the means of support. This shifts the challenge to developing robust numerical methods for removing the influence of the launch pad from the integrated system modal test. The Space Launch System (SLS) is no exception where the mobile launcher (ML) is used to support the vehicle for the integrated modal test (IMT). For the IMT, it is well understood from pre-test analysis with finite element models (FEMs) of the SLS and SLS coupled to ML that the ML has a significant influence on the SLS modal properties especially in the lower frequency range where the primary SLS bending modes exist. An accelerance decoupling (AD) method has been formulated for the purpose of “subtracting out” the influence of the ML from the IMT results. With AD, the SLS decoupled frequency response functions (FRFs) are directly extracted from the IMT FRFs. The subject approach is aimed to utilize measured data only and achieve a robust FRF decoupling scheme. AD is derived from a widely used coupling technique called “Receptance Coupling” (RC). The AD core equation reverses the RC process and utilizes a pair of auxiliary equations that enable the core equation to be resolved based on measured data only. In AD, the decoupled component FRFs are extracted from the coupled system FRFs with a transformation to remove the contribution of the “subtractive component”. This paper addresses the AD’s operational flexibility to resolve SLS free-free modal properties from coupled system measured data but also the possibility to include data from FEM if there is enough confidence in the FEM or if it is asserted that the effect to the final outcome is reasonable.

Accelerance decoupling↗
NASA NTRS2022

Accelerance Decoupling: An Approach for Removing the Influence of the Test Stand from the Integrated Modal Test

The main objective for launch vehicle (LV) modal testing is to quantify the LV’s modal properties in the free-free state (post pad separation). However, given the size of most LV systems, free-free testing is a challenge and often not feasible. With this, a test stand, typically the launch pad itself, is introduced as the means of support. This shifts the challenge to developing robust numerical methods for removing the influence of the launch pad from the integrated system modal test. The Space Launch System (SLS) is no exception where the mobile launcher (ML) is used to support the vehicle for the integrated modal test (IMT). For the IMT, it is well understood from pre-test analysis with finite element models (FEMs) of the SLS and SLS coupled to ML that the ML has a significant influence on the SLS modal properties especially in the lower frequency range where the primary SLS bending modes exist. An accelerance decoupling (AD) method has been formulated for the purpose of “subtracting out” the influence of the ML from the IMT results. With AD, the SLS decoupled frequency response functions (FRFs) are directly extracted from the IMT FRFs. The subject approach is aimed to utilize measured data only and achieve a robust FRF decoupling scheme. AD is derived from a widely used coupling technique called “Receptance Coupling” (RC). The AD core equation reverses the RC process and utilizes a pair of auxiliary equations that enable the core equation to be resolved based on measured data only. In AD, the decoupled component FRFs are extracted from the coupled system FRFs with a transformation to remove the contribution of the “subtractive component”. This paper addresses the AD’s operational flexibility to resolve SLS free-free modal properties from coupled system measured data but also the possibility to include data from FEM if there is enough confidence in the FEM or if it is asserted that the effect to the final outcome is reasonable.

Accelerance decoupling↗
NASA NTRSDate not supplied

Unsteady CFD Simulations of a Compression Corner Geometry Using Wall-Modeled LES Methods in Loci/CHEM

As computing technology continues to improve, simulating unsteady aerodynamic environments on launch vehicle geometries becomes more viable. For NASA’s Space Launch System, the flowfield is characterized by large regions of unsteady turbulent wall-bounded flow. Simulating these unsteady wall-bounded environments is of interest to the SLS Aerodynamics team. Wall-modeled LES methods are currently the best approach for simulating these environments in engineering applications. Less-expensive computationally than fully-resolved methods while maintaining many of the benefits in simulating unsteady environments. Before applying these methods blindly to large geometries, best practices can first be applied and understood on smaller canonical problems.

Bryson Frank↗
NASA NTRS2010

Modified Universal Design Survey: Enhancing Operability of Launch Vehicle Ground Crew Worksites

Operability is a driving requirement for next generation space launch vehicles. Launch site ground operations include numerous operator tasks to prepare the vehicle for launch or to perform preflight maintenance. Ensuring that components requiring operator interaction at the launch site are designed for optimal human use is a high priority for operability. To promote operability, a Design Quality Evaluation Survey based on Universal Design framework was developed to support Human Factors Engineering (HFE) evaluation for NASA s launch vehicles. Universal Design per se is not a priority for launch vehicle processing however; applying principles of Universal Design will increase the probability of an error free and efficient design which promotes operability. The Design Quality Evaluation Survey incorporates and tailors the seven Universal Design Principles and adds new measures for Safety and Efficiency. Adapting an approach proven to measure Universal Design Performance in Product, each principle is associated with multiple performance measures which are rated with the degree to which the statement is true. The Design Quality Evaluation Survey was employed for several launch vehicle ground processing worksite analyses. The tool was found to be most useful for comparative judgments as opposed to an assessment of a single design option. It provided a useful piece of additional data when assessing possible operator interfaces or worksites for operability.

Blume, Jennifer L.↗
NASA NTRS2017

Application of GPS to Enable Launch Vehicle Upper Stage Heliocentric Disposal

To properly dispose of the upper stage of the Space Launch System, the vehicle must perform a burn in Earth orbit to perform a close flyby of the Lunar surface to gain adequate energy to enter into heliocentric space. This architecture was selected to meet NASA requirements to limit orbital debris in the Earth-Moon system. The choice of a flyby for heliocentric disposal was driven by mission and vehicle constraints. This paper describes the SLS mission for Exploration Mission -1, a high level overview of the Block 1 vehicle, and the various disposal options considered. The research focuses on this analysis in terms of the mission design and navigation problem, focusing on the vehicle-level requirements that enable a successful mission. An inertial-only system is shown to be insufficient for heliocentric flyby due to large inertial integration errors from launch through disposal maneuver while on a trans-lunar trajectory. The various options for aiding the navigation system are presented and details are provided on the use of GPS to bound the state errors in orbit to improve the capability for stage disposal. The state estimation algorithm used is described as well as its capability in determination of the vehicle state at the start of the planned maneuver. This data, both dispersions on state and on errors, is then used to develop orbital targets to use for meeting the required Lunar flyby for entering onto a heliocentric trajectory. The effect of guidance and navigation errors on this capability is described as well as the identified constraints for achieving the disposal requirements. Additionally, discussion is provided on continued analysis and identification of system considerations that can drive the ability to integrate onto a vehicle intended for deep space.

Anzalone, Evan J.↗