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NASA NTRS2015

Dynamic Modeling of Ascent Abort Scenarios for Crewed Launches

For the last 30 years, the United States's human space program has been focused on low Earth orbit exploration and operations with the Space Shuttle and International Space Station programs. After nearly 50 years, the U.S. is again working to return humans beyond Earth orbit. To do so, NASA is developing a new launch vehicle and spacecraft to provide this capability. The launch vehicle is referred to as the Space Launch System (SLS) and the spacecraft is called Orion. The new launch system is being developed with an abort system that will enable the crew to escape launch failures that would otherwise be catastrophic as well as probabilistic design requirements set for probability of loss of crew (LOC) and loss of mission (LOM). In order to optimize the risk associated with designing this new launch system, as well as verifying the associated requirements, NASA has developed a comprehensive Probabilistic Risk Assessment (PRA) of the integrated ascent phase of the mission that includes the launch vehicle, spacecraft and ground launch facilities. Given the dynamic nature of rocket launches and the potential for things to go wrong, developing a PRA to assess the risk can be a very challenging effort. Prior to launch and after the crew has boarded the spacecraft, the risk exposure time can be on the order of three hours. During this time, events may initiate from either of the spacecraft, the launch vehicle, or the ground systems, thus requiring an emergency egress from the spacecraft to a safe ground location or a pad abort via the spacecraft's launch abort system. Following launch, again either the spacecraft or the launch vehicle can initiate the need for the crew to abort the mission and return to the home. Obviously, there are thousands of scenarios whose outcome depends on when the abort is initiated during ascent as to how the abort is performed. This includes modeling the risk associated with explosions and benign system failures that require aborting a spacecraft under very dynamic conditions, particularly in the lower atmosphere, and returning the crew home safely. This paper will provide an overview of the PRA model that has been developed of this new launch system, including some of the challenges that are associated with this effort. Key Words: PRA, space launches, human space program, ascent abort, spacecraft, launch vehicles

Bigler, Mark↗
NASA NTRS2015

Dynamic Modeling of Ascent Abort Scenarios for Crewed Launches

For the last 30 years, the United States' human space program has been focused on low Earth orbit exploration and operations with the Space Shuttle and International Space Station programs. After over 40 years, the U.S. is again working to return humans beyond Earth orbit. To do so, NASA is developing a new launch vehicle and spacecraft to provide this capability. The launch vehicle is referred to as the Space Launch System (SLS) and the spacecraft is called Orion. The new launch system is being developed with an abort system that will enable the crew to escape launch failures that would otherwise be catastrophic as well as probabilistic design requirements set for probability of loss of crew (LOC) and loss of mission (LOM). In order to optimize the risk associated with designing this new launch system, as well as verifying the associated requirements, NASA has developed a comprehensive Probabilistic Risk Assessment (PRA) of the integrated ascent phase of the mission that includes the launch vehicle, spacecraft and ground launch facilities. Given the dynamic nature of rocket launches and the potential for things to go wrong, developing a PRA to assess the risk can be a very challenging effort. Prior to launch and after the crew has boarded the spacecraft, the risk exposure time can be on the order of three hours. During this time, events may initiate from either the spacecraft, the launch vehicle, or the ground systems, thus requiring an emergency egress from the spacecraft to a safe ground location or a pad abort via the spacecraft's launch abort system. Following launch, again either the spacecraft or the launch vehicle can initiate the need for the crew to abort the mission and return home. Obviously, there are thousands of scenarios whose outcome depends on when the abort is initiated during ascent and how the abort is performed. This includes modeling the risk associated with explosions and benign system failures that require aborting a spacecraft under very dynamic conditions, particularly in the lower atmosphere, and returning the crew home safely. This paper will provide an overview of the PRA model that has been developed of this new launch system, including some of the challenges that are associated with this effort.

Bigler, Mark↗
NASA NTRS2010

Orion Abort Flight Test

The purpose of NASA's Constellation project is to create the new generation of spacecraft for human flight to the International Space Station in low-earth orbit, the lunar surface, as well as for use in future deep-space exploration. One portion of the Constellation program was the development of the Orion crew exploration vehicle (CEV) to be used in spaceflight. The Orion spacecraft consists of a crew module, service module, space adapter and launch abort system. The crew module was designed to hold as many as six crew members. The Orion crew exploration vehicle is similar in design to the Apollo space capsules, although larger and more massive. The Flight Test Office is the responsible flight test organization for the launch abort system on the Orion crew exploration vehicle. The Flight Test Office originally proposed six tests that would demonstrate the use of the launch abort system. These flight tests were to be performed at the White Sands Missile Range in New Mexico and were similar in nature to the Apollo Little Joe II tests performed in the 1960s. The first flight test of the launch abort system was a pad abort (PA-1), that took place on 6 May 2010 at the White Sands Missile Range in New Mexico. Primary flight test objectives were to demonstrate the capability of the launch abort system to propel the crew module a safe distance away from a launch vehicle during a pad abort, to demonstrate the stability and control characteristics of the vehicle, and to determine the performance of the motors contained within the launch abort system. The focus of the PA-1 flight test was engineering development and data acquisition, not certification. In this presentation, a high level overview of the PA-1 vehicle is given, along with an overview of the Mobile Operations Facility and information on the White Sands tracking sites for radar & optics. Several lessons learned are presented, including detailed information on the lessons learned in the development of wind placards for flight. PA-1 flight data is shown, as well as a comparison of PA-1 flight data to nonlinear simulation Monte Carlo data.

Hayes, Peggy Sue↗
NASA NTRS2011

Launch Vehicle Demonstrator Using Shuttle Assets

The Marshall Space Flight Center Advanced Concepts Office (ACO) has the leading role for NASA s preliminary conceptual launch vehicle design and performance analysis. Over the past several years the ACO Earth-to-Orbit Team has evaluated thousands of launch vehicle concept variations for a multitude of studies including agency-wide efforts such as the Exploration Systems Architecture Study (ESAS), Constellation, Heavy Lift Launch Vehicle (HLLV), Heavy Lift Propulsion Technology (HLPT), Human Exploration Framework Team (HEFT), and Space Launch System (SLS). NASA plans to continue human space exploration and space station utilization. Launch vehicles used for heavy lift cargo and crew will be needed. One of the current leading concepts for future heavy lift capability is an inline one and a half stage concept using solid rocket boosters (SRB) and based on current Shuttle technology and elements. Potentially, the quickest and most cost-effective path towards an operational vehicle of this configuration is to make use of a demonstrator vehicle fabricated from existing shuttle assets and relying upon the existing STS launch infrastructure. Such a demonstrator would yield valuable proof-of-concept data and would provide a working test platform allowing for validated systems integration. Using shuttle hardware such as existing RS-25D engines and partial MPS, propellant tanks derived from the External Tank (ET) design and tooling, and four-segment SRB s could reduce the associated upfront development costs and schedule when compared to a concept that would rely on new propulsion technology and engine designs. There are potentially several other additional benefits to this demonstrator concept. Since a concept of this type would be based on man-rated flight proven hardware components, this demonstrator has the potential to evolve into the first iteration of heavy lift crew or cargo and serve as a baseline for block upgrades. This vehicle could also serve as a demonstration and test platform for the Orion Program. Critical spacecraft systems, re-entry and recovery systems, and launch abort systems of Orion could also be demonstrated in early test flights of the launch vehicle demo. Furthermore, an early demonstrator of this type would provide a stop-gap for retaining critical human capital and infrastructure while affording the current emerging generation of young engineers opportunity to work with and capture lessons learned from existing STS program offices and personnel, who were integral in the design and development of the Space Shuttle before these resources are no longer available. The objective of this study is to define candidate launch vehicle demonstration concepts that are based on Space Shuttle assets and determine their performance capabilities and how these demonstration vehicles could evolve to a heavy lift capability to low earth orbit.

Threet, Grady E., Jr.↗
NASA NTRS2011

MLAS and LAS: Project Comparison of Lessons Learned

The development of new and safer manned space vehicles is a top priority at NASA. Recently two different approaches of how to accomplish this mission of keeping astronauts safe was successfully demonstrated. With work already underway on an Apollo-like launch abort system for the Orion Crew Exploration Vehicle (CEV), an alternative design concept named the Max Launch Abort System, or MLAS, was developed as a parallel effort. The Orion system, managed by the Constellation office, is based on the design of a single solid launch abort motor in a tower positioned above the capsule. The MLAS design takes a different approach placing the solid launch abort motor underneath the capsule. This effort was led by the NASA Engineering and Safety Center (NESC). Both escape systems were designed with the Ares I Rocket as the launch vehicle and had the same primary requirement to safely propel a crew module away from any emergency event either on the launch pad or during accent. Beyond these two parameters, there was little else in common between the two projects, except that they both concluded in successful launches that will further promote the development of crew launch abort systems. This presentation will compare both projects from the standpoint of technical requirements; program management and flight test objectives and highlight the synergistic lessons learned by two engineers who worked on each program.

Dittemore, Gary↗
NASA NTRS2011

Comparison of Two Recent Launch Abort Platforms

The development of new and safer manned space vehicles is a top priority at NASA. Recently two different approaches of how to accomplish this mission of keeping astronauts safe was successfully demonstrated. With work already underway on an Apollo-like launch abort system for the Orion Crew Exploration Vehicle (CEV), an alternative design concept named the Max Launch Abort System, or MLAS, was developed as a parallel effort. The Orion system, managed by the Constellation office, is based on the design of a single solid launch abort motor in a tower positioned above the capsule. The MLAS design takes a different approach placing the solid launch abort motor underneath the capsule. This effort was led by the NASA Engineering and Safety Center (NESC). Both escape systems were designed with the Ares I Rocket as the launch vehicle and had the same primary requirement to safely propel a crew module away from any emergency event either on the launch pad or during accent. Beyond these two parameters, there was little else in common between the two projects, except that they both concluded in successful launches that will further promote the development of crew launch abort systems. A comparison of these projects from the standpoint of technical requirements; program management and flight test objectives will be done to highlight the synergistic lessons learned by two engineers who worked on each program. This comparison will demonstrate how the scope of the project architecture and management involvement in innovation should be tailored to meet the specific needs of the system under development.

Dittemore, Gary D.↗
NASA NTRSDate not supplied

Artemis Innovative Assembly and Integration Operations of the Launch Abort Systemat KSC

This paper describes the Artemis program assembly and integration approach that enables affordable and low risk processing operations for the Launch Abort System at KSC. NASA is currently developing the Artemis 1 spacecraft to meet the test objectives of an uncrewed orbital mission to the moon and return to earth in 2021. The Artemis Orion spacecraft consists of a Crew Module (CM), a Service Module (SM), and a Launch Abort System (LAS) to safely transport a crew of 4 to the moon and return to earth. The LAS enables the CM to have an abort capability for the crew during the launch phase of the mission from the launch pad throughout the ascent phase. The LAS includes motor elements for abort, attitude control, and stage separation is assembled at the Launch Abort System Facility (LASF) adjacent to the CM/SM (CSM) assembly operations in the Neil Armstrong Operations and Checkout Building (O&C) at KSC. The LAS integration to the CSM is completed in the LASF and the completed Artemis spacecraft is integrated to the Space Launch System (SLS) in the Vehicle Assembly Building (VAB) at KSC. A significant LAS development flight test milestone was recently completed in 2019 for the Ascent Abort (AA-2) flight test at Cape Canaveral Air Force Station where a high-altitude abort test was successfully achieved using a Peacekeeper booster stage exercising all of the LAS systems including propulsion, avionics, and pyrotechnics demonstrating he LAS readiness for crewed flight. The first operational flight of the LAS will be for the Artemis 2 mission which is the first crewed mission for the Orion program planned in 2023.

William J Koenig↗
NASA NTRS2009

A Review of the MLAS Parachute Systems

The NASA Engineering and Safety Center (NESC) is developing the Max Launch Abort System (MLAS) as a risk-mitigation design should problems arise with the baseline Orion spacecraft launch abort design. The Max in MLAS is dedicated to Max Faget, the renowned NASA spacecraft designer. The MLAS flight test vehicle consists of boost skirt, coast skirt and the MLAS fairing which houses a full scale boilerplate Orion Crew Module (CM). The objective of the flight test is to prove that the CM can be released from the MLAS fairing during pad abort conditions without detrimental recontact between the CM and fairing, achieving performance similar to the Orion launch abort system. The boost and coast skirts provide the necessary thrust and stability to achieve the flight test conditions and are released prior to the test -- much like the Little Joe booster was used in the Apollo Launch Escape System tests. To achieve the test objective, two parachutes are deployed from the fairing to reorient the CM/fairing to a heatshield first orientation. The parachutes then provide the force necessary to reduce the total angle of attack and body angular rates required for safe release of the CM from the fairing. A secondary test objective after CM release from the fairing is to investigate the removal of the CM forward bay cover (FBC) with CM drogue parachutes for the purpose of attempting to synchronously deploying a set of CM main parachutes. Although multiple parachute deployments are used in the MLAS flight test vehicle to complete its objective, there are only two parachute types employed in the flight test. Five of the nine parachutes used for MLAS are 27.6 ft D(sub 0) ribbon parachutes, and the remaining four are standard G-12 cargo parachutes. This paper presents an overview of the 27.6 ft D(sub 0) ribbon parachute system employed on the MLAS flight test vehicle for coast skirt separation, fairing reorientation, and as drogue parachutes for the CM after separation from the fairing. Discussion will include: the process used to select this design, previously proven as a spin/stall recovery parachute; descriptions of all components of the parachute system; the minor modifications necessary to adapt the parachute to the MLAS program; the techniques used to analyze the parachute for the multiple roles it performs; a discussion of the rigging techniques used to interface the parachute system to the vehicle; and a brief description of how the evolution of the program affected parachute usage and analysis. An overview of the Objective system, rationale for the MLAS approach and the future of the program will also be presented. We hope to have flight test results to report at the time of the Conference Presentation.

Taylor, Anthony P.↗
NASA NTRSDate not supplied

Safety & Mission Assurance (SMA) Activities in Support of Artemis I and Plans for Artemis II

The Artemis I mission in 2022 accomplished humanity’s first venture beyond Low-Earth Orbit (LEO) with a human-rated spacecraft in over fifty years. Artemis comprises several key Program elements – Multi-Purpose Crew Vehicle (MPCV) Orion Crew Module, Service Module, Launch Abort System; the Space Launch System (SLS) and the Exploration Ground System (EGS). Much of our human spaceflight experience with crewed lunar missions is captured in historical program documents, but first-hand knowledge is limited to a few spaceflight veterans. Missions to LEO have offered the opportunity of direct, near-instantaneous communications and assistance, and the ability to return to Earth within a matter of hours in case of emergency. Outward-bound missions do not have these features and will require a more autonomous and reliable spacecraft. The assessment of compliance with applicable safety requirements and adequacy of hazard controls and verifications is the responsibility of the MPCV Safety and Engineering Review Panel (MSERP), along with a Joint [NASA/ESA] Safety and Engineering Review Panel (JSERP) for the review of the European Service Module (ESM). The MSERP has two features that are relatively unique amongst NASA safety panels. First, NASA Engineering was added as a Panel co-chair, which enhanced the Engineering organization’s engagement and level of understanding of hazard analysis methodology and results. The MSERP has representation from each office, including Flight Operations Directorate (FOD) and Health and Medical Technical Authority, supporting the Orion Program in addition to the Engineering and Safety Technical Authority co-chairs. Second, the JSERP has a second set of co-chairs from ESA Engineering and ESA Product Assurance and Safety organizations, which recognizes the international arrangement as one of partnership. This paper will focus on SMA processes, activities, and plans for the Orion element and explore unique challenges associated with Artemis II as we approach the flight of the first crewed Orion vehicle.

Paul J. Collier↗
NASA NTRS2009

A Flight Dynamics Perspective of the Orion Pad Abort One Flight Test

The Orion Crew Exploration Vehicle is America s next generation of human rated spacecraft. The Orion Launch Abort System will take the astronauts away from the exploration vehicle in the event of an aborted launch. The pad abort mode of the Launch Abort System will be flight-tested in 2009 from the White Sands Missile Range in New Mexico. This paper examines some of the efforts currently underway at the NASA Dryden Flight Research Center by the Controls & Dynamics group in preparation for the flight test. The concept of operation for the pad abort flight is presented along with an overview of the guidance, control and navigation systems. Preparations for the flight test, such as hardware testing and development of the real-time displays, are examined. The results from the validation and verification efforts for the aerodynamic and atmospheric models are shown along with Monte Carlo analysis results.

Idicula, Jinu↗
NASA NTRS2009

Ares I-X: On the Threshold of Exploration

Ares I-X, the first flight of the Ares I crew launch vehicle, is less than a year from launch. Ares I-X will test the flight characteristics of Ares I from liftoff to first stage separation and recovery. The flight also will demonstrate the computer hardware and software (avionics) needed to control the vehicle; deploy the parachutes that allow the first stage booster to land in the ocean safely; measure and control how much the rocket rolls during flight; test and measure the effects of first stage separation; and develop and try out new ground handling and rocket stacking procedures in the Vehicle Assembly Building (VAB) and first stage recovery procedures at Kennedy Space Center (KSC) in Florida. All Ares I-X major elements have completed their critical design reviews, and are nearing final fabrication. The first stage--four-segment solid rocket booster from the Space Shuttle inventory--incorporates new simulated forward structures to match the Ares I five-segment booster. The upper stage, Orion crew module, and launch abort system will comprise simulator hardware that incorporates developmental flight instrumentation for essential data collection during the mission. The upper stage simulator consists of smaller cylindrical segments, which were transported to KSC in fall 2008. The crew module and launch abort system simulator were shipped in December 2008. The first stage hardware, active roll control system (RoCS), and avionics components will be delivered to KSC in 2009. This paper will provide detailed statuses of the Ares I-X hardware elements as NASA's Constellation Program prepares for this first flight of a new exploration era in the summer of 2009.

Davis, Stephan R.↗
NASA NTRS2013

Orion MPCV GN and C End-to-End Phasing Tests

End-to-end integration tests are critical risk reduction efforts for any complex vehicle. Phasing tests are an end-to-end integrated test that validates system directional phasing (polarity) from sensor measurement through software algorithms to end effector response. Phasing tests are typically performed on a fully integrated and assembled flight vehicle where sensors are stimulated by moving the vehicle and the effectors are observed for proper polarity. Orion Multi-Purpose Crew Vehicle (MPCV) Pad Abort 1 (PA-1) Phasing Test was conducted from inertial measurement to Launch Abort System (LAS). Orion Exploration Flight Test 1 (EFT-1) has two end-to-end phasing tests planned. The first test from inertial measurement to Crew Module (CM) reaction control system thrusters uses navigation and flight control system software algorithms to process commands. The second test from inertial measurement to CM S-Band Phased Array Antenna (PAA) uses navigation and communication system software algorithms to process commands. Future Orion flights include Ascent Abort Flight Test 2 (AA-2) and Exploration Mission 1 (EM-1). These flights will include additional or updated sensors, software algorithms and effectors. This paper will explore the implementation of end-to-end phasing tests on a flight vehicle which has many constraints, trade-offs and compromises. Orion PA-1 Phasing Test was conducted at White Sands Missile Range (WSMR) from March 4-6, 2010. This test decreased the risk of mission failure by demonstrating proper flight control system polarity. Demonstration was achieved by stimulating the primary navigation sensor, processing sensor data to commands and viewing propulsion response. PA-1 primary navigation sensor was a Space Integrated Inertial Navigation System (INS) and Global Positioning System (GPS) (SIGI) which has onboard processing, INS (3 accelerometers and 3 rate gyros) and no GPS receiver. SIGI data was processed by GN&C software into thrust magnitude and direction commands. The processing changes through three phases of powered flight: pitchover, downrange and reorientation. The primary inputs to GN&C are attitude position, attitude rates, angle of attack (AOA) and angle of sideslip (AOS). Pitch and yaw attitude and attitude rate responses were verified by using a flight spare SIGI mounted to a 2-axis rate table. AOA and AOS responses were verified by using a data recorded from SIGI movements on a robotic arm located at NASA Johnson Space Center. The data was consolidated and used in an open-loop data input to the SIGI. Propulsion was the Launch Abort System (LAS) Attitude Control Motor (ACM) which consisted of a solid motor with 8 nozzles. Each nozzle has active thrust control by varying throat area with a pintle. LAS ACM pintles are observable through optically transparent nozzle covers. SIGI movements on robot arm, SIGI rate table movements and LAS ACM pintle responses were video recorded as test artifacts for analysis and evaluation. The PA-1 Phasing Test design was determined based on test performance requirements, operational restrictions and EGSE capabilities. This development progressed during different stages. For convenience these development stages are initial, working group, tiger team, Engineering Review Team (ERT) and final.

Neumann, Brian C.↗
NASA NTRS2010

Ares I-X Flight Test Vehicle: Stack 5 Modal Test

Ares I-X was the first flight test vehicle used in the development of NASA's Ares I crew launch vehicle. The Ares I-X used a 4-segment reusable solid rocket booster from the Space Shuttle heritage with mass simulators for the 5th segment, upper stage, crew module and launch abort system. Three modal tests were defined to verify the dynamic finite element model of the Ares I-X flight test vehicle. Test configurations included two partial stacks and the full Ares I-X flight test vehicle on the Mobile Launcher Platform. This report focuses on the first modal test that was performed on the top section of the vehicle referred to as Stack 5, which consisted of the spacecraft adapter, service module, crew module and launch abort system simulators. This report describes the test requirements, constraints, pre-test analysis, test operations and data analysis for the Ares I-X Stack 5 modal test.

Buehrle, Ralph D.↗
NASA NTRS2015

The Orion Pad Abort 1 (PA-1) Flight Test: A Propulsion Success

This poster provides a concise overview of the highly successful Orion Pad Abort 1 (PA-1) flight test, and the three rocket motors that contributed to this success. The primary purpose of the Orion PA-1 flight was to help certify the Orion Launch Abort System (LAS), which can be utilized in the unlikely event of an emergency on the launchpad or during mission vehicle ascent. The PA-1 test was the first fully integrated flight test of the Orion LAS, one of the primary systems within the Orion Multi-Purpose Crew Vehicle (MPCV). The Orion MPCV is part of the architecture within the Space Launch System (SLS), which is being designed to transport astronauts beyond low-Earth orbit for future exploration missions. Had the Orion PA-1 flight abort occurred during launch preparations for a real human spaceflight mission, the PA-1 LAS would have saved the lives of the crew. The PA-1 flight test was largely successful due to the three solid rocket motors of the LAS: the Attitude Control Motor (ACM); the Jettison Motor (JM); and the Abort Motor (AM). All three rocket motors successfully performed their required functions during the Orion PA-1 flight test, flown on May 6, 2010 at the White Sands Missile Range in New Mexico, culminating in a successful demonstration of an abort capability from the launchpad.

JM↗
NASA NTRS2011

First Stage Solid Propellant Multiply Debris Thermal Analysis

Destruction of a solid rocket stage of a launch vehicle can create a thermal radiation hazard for an aborting crew module. This hazard was assessed for the Constellation Program (Cx) crew and launch vehicle concept. For this concept, if an abort was initiated in first stage flight, the Crew Module (CM) will separate and be pulled away from the malfunctioning launch vehicle via a Launch Abort System (LAS). Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment, caused by surrounding burning propellant debris, may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and to quantify the risk of first stage propellant debris leading to the thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Toleman, Benjamin M.↗
NASA NTRS2016

Overview of the Space Launch System Ascent Aeroacoustic Environment Test Program

Characterization of accurate flight vehicle unsteady aerodynamics is critical for component and secondary structure vibroacoustic design. The Aerosciences Branch at the National Aeronautics and Space Administration (NASA) Marshall Space Flight Center has conducted a test at the NASA Ames Research Center (ARC) Unitary Plan Wind Tunnels (UPWT) to determine such ascent aeroacoustic environments for the Space Launch System (SLS). Surface static pressure measurements were also collected to aid in determination of local environments for venting, CFD substantiation, and calibration of the flush air data system located on the launch abort system. Additionally, this test supported a NASA Engineering and Safety Center study of alternate booster nose caps. Testing occurred during two test campaigns: August - September 2013 and December 2013 - January 2014. Four primary model configurations were tested for ascent aeroacoustic environment definition. The SLS Block 1 vehicle was represented by a 2.5% full stack model and a 4% truncated model. Preliminary Block 1B payload and manned configurations were also tested, using 2.5% full stack and 4% truncated models respectively. This test utilized the 11 x 11 foot transonic and 9 x 7 foot supersonic tunnel sections at the ARC UPWT to collect data from Mach 0.7 through 2.5 at various total angles of attack. SLS Block 1 design environments were developed primarily using these data. SLS Block 1B preliminary environments have also been prepared using these data. This paper discusses the test and analysis methodology utilized, with a focus on the unsteady data collection and processing.

Herron, Andrew J.↗
NASA NTRS2011

First Stage Solid Propellant Multi Debris Thermal Analysis

The crew launch vehicle considered for the Constellation (Cx) Program utilizes a first stage solid rocket motor. If an abort is initiated in first stage flight the Crew Module (CM) will separate and be pulled away from the launch vehicle via a Launch Abort System (LAS) in order to safely and quickly carry the crew away from the malfunction launch vehicle. Having aborted the mission, the launch vehicle will likely be destroyed via a Flight Termination System (FTS) in order to prevent it from errantly traversing back over land and posing a risk to the public. The resulting launch vehicle debris field, composed primarily of first stage solid propellant, poses a threat to the CM. The harsh radiative thermal environment induced by surrounding burning propellant debris may lead to CM parachute failure. A methodology, detailed herein, has been developed to address this concern and quantify the risk of first stage propellant debris leading to radiative thermal demise of the CM parachutes. Utilizing basic thermal radiation principles, a software program was developed to calculate parachute temperature as a function of time for a given abort trajectory and debris piece trajectory set. Two test cases, considered worst-case aborts with regard to launch vehicle debris environments, were analyzed using the simulation: an abort declared at Mach 1 and an abort declared at maximum dynamic pressure (Max Q). For both cases, the resulting temperature profiles indicated that thermal limits for the parachutes were not exceeded. However, short duration close encounters by single debris pieces did have a significant effect on parachute temperature, with magnitudes on the order of 10 s of degrees Fahrenheit. Therefore while these two test cases did not indicate exceedance of thermal limits, in order to quantify the risk of parachute failure due to radiative effects from the abort environment, a more thorough probability-based analysis using the methodology demonstrated herein must be performed.

Toleman, Benjamin M.↗
NASA NTRS2010

The Range Safety Debris Catalog Analysis in Preparation for the Pad Abort One Flight Test

The Pad Abort One flight test of the Orion Abort Flight Test Program is currently under development with the goal of demonstrating the capability of the Launch Abort System. In the event of a launch failure, this system will propel the Crew Exploration Vehicle to safety. An essential component of this flight test is range safety, which ensures the security of range assets and personnel. A debris catalog analysis was done as part of a range safety data package delivered to the White Sands Missile Range in New Mexico where the test will be conducted. The analysis discusses the consequences of an overpressurization of the Abort Motor. The resulting structural failure was assumed to create a debris field of vehicle fragments that could potentially pose a hazard to the range. A statistical model was used to assemble the debris catalog of potential propellant fragments. Then, a thermodynamic, energy balance model was applied to the system in order to determine the imparted velocity to these propellant fragments. This analysis was conducted at four points along the flight trajectory to better understand the failure consequences over the entire flight. The methods used to perform this analysis are outlined in detail and the corresponding results are presented and discussed.

Kutty, Prasad M.↗