Results for “launch separation” — Page 27

NASA NTRS2010

Mission and Implementation of an Affordable Lunar Return

We present an architecture that establishes the infrastructure for routine space travel by taking advantage of the Moon's resources, proximity and accessibility. We use robotic assets on the Moon that are teleoperated from Earth to prospect, test, demonstrate and produce water from lunar resources before human arrival. This plan is affordable, flexible and not tied to any specific launch vehicle solution. Individual surface pieces are small, permitting them to be deployed separately on small launchers or combined together on single large launchers. Schedule is our free variable; even under highly constrained budgets, the architecture permits this program to be continuously pursued using small, incremental, cumulative steps. The end stage is a fully functional, human-tended lunar outpost capable of producing 150 metric tonnes of water per year enough to export water from the Moon and create a transportation system that allows routine access to all of cislunar space. This cost-effective lunar architecture advances technology and builds a sustainable transportation infrastructure. By eliminating the need to launch everything from the surface of the Earth, we fundamentally change the paradigm of spaceflight.

Spudis, Paul↗

NASA NTRS2017

NASA Collaborative Design Processes

This is Block 1, the first evolution of the world's most powerful and versatile rocket, the Space Launch System, built to return humans to the area around the moon. Eventually, larger and even more powerful and capable configurations will take astronauts and cargo to Mars. On the sides of the rocket are the twin solid rocket boosters that provide more than 75 percent during liftoff and burn for about two minutes, after which they are jettisoned, lightening the load for the rest of the space flight. Four RS-25 main engines provide thrust for the first stage of the rocket. These are the world's most reliable rocket engines. The core stage is the main body of the rocket and houses the fuel for the RS-25 engines, liquid hydrogen and liquid oxygen, and the avionics, or "brain" of the rocket. The core stage is all new and being manufactured at NASA's "rocket factory," Michoud Assembly Facility near New Orleans. The Launch Vehicle Stage Adapter, or LVSA, connects the core stage to the Interim Cryogenic Propulsion Stage. The Interim Cryogenic Propulsion Stage, or ICPS, uses one RL-10 rocket engine and will propel the Orion spacecraft on its deep-space journey after first-stage separation. Finally, the Orion human-rated spacecraft sits atop the massive Saturn V-sized launch vehicle. Managed out of Johnson Space Center in Houston, Orion is the first spacecraft in history capable of taking humans to multiple destinations within deep space. 2) Each element of the SLS utilizes collaborative design processes to achieve the incredible goal of sending human into deep space. Early phases are focused on feasibility and requirements development. Later phases are focused on detailed design, testing, and operations. There are 4 basic phases typically found in each phase of development.

Jones, Davey↗

NASA NTRSDate not supplied

Risk Reduction Project in Pad Abort-1 Launch Vehicle Loads & Dynamics

The Pad Abort (PA-1) system includes the Launch Abort System (LAS) and the Command Module (CM). The PA-1 abort flight test will launch from the White Sands Missile Range. Prior to the ignition, the vehicle will be resting, without restraints, on a short support structure. The Launch Abort Motor will ignite and burn for less than 5 seconds before a 20 second coast. Then the LAS will separate from the CM and the CM will descend via parachute to a soil landing. The static firing of the Abort Motor (ST-1) for PA-1 Launch Vehicle resulted in unexpected higher levels that superseded the environment predictions and all the design and test loads of all subassemblies and components. A rapid project was put in place to develop the flight environments, loads and associated uncertainties for the Verification Load Cycle (VLC) of PA-1. An acoustic test and shaker test of the CM were planned and executed in support of the developments of damping and transfer functions. Instead of using the traditional structure-borne envelops, the actual ST-1 pressures and forces were used to develop the internal accelerations. The PA-1 motor was cooled to 60 Deg F using thermal conditioning to reduce the thrust profile. The plumes of the abort motor were studied to determine if the airborne environments would be reduced. The combination of all items, noted above, allowed a reduction in loads for VLC.

Sasan C Armand↗

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 NTRS1995

Atlas/Centaur-SOHO Pre-Launch News Conference

Live coverage of a pre-launch news conference on the Atlas/Centaur SOHO mission is presented. George Diller, NASA Public Affairs, introduces the panel. Floyd Curington, NASA Launch Manager, Kennedy Space Center, presents countdown activities. Pat Symons, Launch Vehicle Manager, NASA Lewis Research Center, analyzes the time duration from liftoff to spacecraft separation. Fabrizio Felici, SOHO Mission Director European Space Agency (ESA), explains the important features of SOHO, which includes a payload of 650 kilos and 12 major instruments with multisensors. Ken Sizemore, International Solar Terrestrial Physics (ISTP) Project Manager Goddard Space Flight Center (GSFC), talks about the successful international collaboration between the ESA and NASA. Joel Tumbiolo, Launch Weather Officer United States Air Force (USAF, presented the weather forecast. SOHO was launched aboard an Atlas II rocket on November 23, 1995. The news conference ends with a brief question and answer period.

Source record↗

NASA NTRS2015

Estimating the Life Cycle Cost of Space Systems

A space system's Life Cycle Cost (LCC) includes design and development, launch and emplacement, and operations and maintenance. Each of these cost factors is usually estimated separately. NASA uses three different parametric models for the design and development cost of crewed space systems; the commercial PRICE-H space hardware cost model, the NASA-Air Force Cost Model (NAFCOM), and the Advanced Missions Cost Model (AMCM). System mass is an important parameter in all three models. System mass also determines the launch and emplacement cost, which directly depends on the cost per kilogram to launch mass to Low Earth Orbit (LEO). The launch and emplacement cost is the cost to launch to LEO the system itself and also the rockets, propellant, and lander needed to emplace it. The ratio of the total launch mass to payload mass depends on the mission scenario and destination. The operations and maintenance costs include any material and spares provided, the ground control crew, and sustaining engineering. The Mission Operations Cost Model (MOCM) estimates these costs as a percentage of the system development cost per year.

Life cycle cost↗

NASA NTRS1991

A liquid/gas phase separator for He-I and He-II

A liquid/gas phase separator has been developed which separates both liquid He-I and He-II from their vapor. The phase separator was designed for the Superfluid Helium On Orbit Transfer (SHOOT) Flight Demonstration both to cool the liquid He after launch (at temperatures between 2.8 and 4.3 K) to the operating temperature of 1.4 K and as a low rate vent on orbit to maintain operating temperature. The phase separator is made of high-purity copper disks held apart by 6 micron Kevlar fibers. It works on the principle of conducting heat from within the dewar to vaporize liquid as it is throttled in the slits to efficiently cool the remaining liquid. Laboratory tests have demonstrated perfect phase separation for both He at its saturated vapor pressure from 1.2 to 4.3 K and for He-II at 2.13 K at pressures from 4.6 to 112 kPa. The performance of this phase separator during lab testing as well as expected performance in space is discussed.

Shirron, P. J.↗

NASA NTRS1971

Saturn S-2 problem resolution history report

A summary of S-2 Program problems and the solutions that were implemented is presented. The problems occurred during a period starting with the initial design concepts and continuing through the launch of the tenth S-2 flight stage information is from nine separate disciplines: design, facilities, logistics, manufacturing, material, program management, quality assurance, safety, and tests.

Virgil, F. W.↗

NASA NTRS1972

Experiments on the dynamic stability of the space shuttle

Some dynamic stability experiments on the space shuttle using a half-model oscillatory technique are discussed. Resulting information from the experiments is presented and includes: (1) dynamic pitching characteristics of both the orbiter and the booster alone as well as of the two models mated into a single launch configuration; (2) the static and dynamic interference effects during an abort separation maneuver; and (3) the dynamic plume interference effect.

Orlik-Rueckermann, K. J.↗

NASA NTRS1993

Corvid meteoroids are not ejecta from the Giordano Bruno impact

The study points out three difficulties with the hypothesis of Hartung (1993) that the Corvid meteor system, observed only once in 1937, may be the return of the clump of ejecta from the formation of a lunar crater, specifically, an event recorded in the chronicles of Gervase of Canterbury on June 25, 1178, which Hartung (1976) previously suggested may be an eyewitness account of the formation of the lunar crater Giordano Bruno. On the basis of this, he predicts that another Corvid shower may be observed in 2003 or 2006. This hypothesis is rejected on the contention that it is implausible that a clump of ejecta could be launched into heliocentric orbit with a low enough dispersion in velocity among separate pieces that it would produce a meteor shower in just one year and not others. Subsequent perturbations by the earth on parts of the clump passing near the earth but not impacting would destroy the coherence of the clump on a time scale much shorter than the 759-yr interval proposed. There are so many orbits that could yield a shower after 759 yr that it is unlikely that a prediction of a return in a specific year would be correct.

Harris, A. W.↗

NASA NTRS2002

Space Flight Qualification Program for the AMS-2 Commercial Cryocoolers

The Alpha Magnetic Spectrometer-02 (AMS-02) experiment is a state-of-the-art particle physics detector containing a large superfluid helium-cooled superconducting magnet. Highly sensitive detector plates inside the magnet measure a particle's speed, momentum, charge, and path. The AMS-02 experiment will study the properties and origin of cosmic particles and nuclei including antimatter and dark matter. AMS-02 will be installed on the International Space Station on Utilization Flight-4. The experiment will be run for at least three years. To extend the life of the stored cryogen and minimize temperature gradients around the magnet, four Stirling-cycle Sunpower M87N cryocoolers will be integrated with AMS-02. The cryocooler cold tip will be connected via a flexible strap to the outer vapor cooled shield of the dewar. Initial thermal analysis shows the lifetime of the experiment is increased by a factor of 2.8 with the use of the cryocooler. The AMS-02 project selected the Sunpower M87 cryocoolers and has asked NASA Goddard to qualify the cryocoolers for space flight use. This paper describes the interfaces with the cryocoolers and presents data collected during testing of the two engineering model cryocoolers. Tests include thermal performance characterization and launch vibration testing. Magnetic field compatibility testing will be presented in a separate paper at the conference.

Shirey, K. A.↗

NASA NTRS2004

A Summary of the Space Shuttle Columbia Tragedy and the Use of LS Dyna in the Accident Investigation and Return to Flight Efforts

On February 1, 2003, the Space Shuttle Columbia broke apart during reentry resulting in loss of 7 crewmembers and craft. For the next several months an extensive investigation of the accident ensued involving a nationwide team of experts from NASA, industry, and academia, spanning dozens of technical disciplines. The Columbia Accident Investigation Board (CAIB), a group of experts assembled to conduct an investigation independent of NASA concluded in August, 2003 that the cause of the loss of Columbia and its crew was a breach in the left wing leading edge Reinforced Carbon-Carbon (RCC) thermal protection system initiated by the impact of thermal insulating foam that had separated from the orbiters external fuel tank 81 seconds into the missions launch. During reentry, this breach allowed superheated air to penetrate behind the leading edge and erode the aluminum structure of left wing which ultimately led to the breakup of the orbiter. In order to gain a better understanding the foam impact on the orbiters RCC wing leading edge, a multi-center team of NASA and Boeing impact experts was formed to characterize the foam and RCC materials for impact analysis using LS Dyna. Dyna predictions were validated with sub-component and full scale tests. LS Dyna proved to be a valuable asset in supporting both the Columbia Accident Investigation and NASA s return to flight efforts. This paper summarizes Columbia Accident and the nearly seven month long investigation that followed. The use of LS-DYNA in this effort is highlighted. Contributions to the investigation and return to flight efforts of the multicenter team consisting of members from NASA Glenn, NASA Langley, and Boeing Philadelphia are introduced and covered in detail in papers to follow in these proceedings.

Melis, Matthew↗

NASA NTRS2003

Rollable Thin-Shell Nanolaminate Mirrors

A class of lightweight, deployable, thin-shell, curved mirrors with built-in precise-shape-control actuators is being developed for high-resolution scientific imaging. This technology incorporates a combination of advanced design concepts in actuation and membrane optics that, heretofore, have been considered as separate innovations. These mirrors are conceived to be stowed compactly in a launch shroud and transported aboard spacecraft, then deployed in outer space to required precise shapes at much larger dimensions (diameters of the order of meters or tens of meters). A typical shell rollable mirror structure would include: (1) a flexible single- or multiple-layer face sheet that would include an integrated reflective surface layer that would constitute the mirror; (2) structural supports in the form of stiffeners made of a shape-memory alloy (SMA); and (3) piezoelectric actuators. The actuators, together with an electronic control subsystem, would implement a concept of hierarchical distributed control, in which (1) the SMA actuators would be used for global shape control and would generate the large deformations needed for the deployment process and (2) the piezoelectric actuators would generate smaller deformations and would be used primarily to effect fine local control of the shape of the mirror.

Hickey, Gregory↗

NASA NTRS2012

CFD Assessment of Forward Booster Separation Motor Ignition Overpressure on ET XT 718 Ice/Frost Ramp

Computational fluid dynamics assessment of the forward booster separation motor ignition over-pressure was performed on the space shuttle external tank X(sub T) 718 ice/frost ramp using the flow solver OVERFLOW. The main objective of this study was the investigation of the over-pressure during solid rocket booster separation and its affect on the local pressure and air-load environments. Delta pressure and plume impingement were investigated as a possible contributing factor to the cause of the debris loss on shuttle missions STS-125 and STS-127. A simplified computational model of the Space Shuttle Launch Vehicle was developed consisting of just the external tank and the solid rocket boosters with separation motor nozzles and plumes. The simplified model was validated by comparison to full fidelity computational model of the Space Shuttle without the separation motors. Quasi steady-state plume solutions were used to calibrate the thrust of the separation motors. Time-accurate simulations of the firing of the booster-separation motors were performed. Parametric studies of the time-step size and the number of sub-iterations were used to find the best converged solution. The computed solutions were compared to previous OVERFLOW steady-state runs of the separation motors with reaction control system jets and to ground test data. The results indicated that delta pressure from the overpressure was small and within design limits, and thus was unlikely to have contributed to the foam losses.

Tejnil, Edward↗

NASA NTRS2016

Evolution of Orion Mission Design for Exploration Mission 1 and 2

The evolving mission design and concepts of NASA's next steps have shaped Orion into the spacecraft that it is today. Since the initial inception of Orion, through the Constellation Program, and now in the Exploration Mission frame-work with the Space Launch System (SLS), each mission design concept and program goal have left Orion with a set of capabilities that can be utilized in many different mission types. Exploration Missions 1 and 2 (EM-1 and EM-2) have now been at the forefront of the mission design focus for the last several years. During that time, different Design Reference Missions (DRMs) were built, analyzed, and modified to solve or mitigate enterprise level design trades to ensure a viable mission from launch to landing. The resulting DRMs for EM-1 and EM-2 were then expanded into multi-year trajectory scans to characterize vehicle performance and Earth-Moon geometry trends. This provides Orion's subsystems with stressing reference trajectories to help design their system. Now that Orion has progressed through the Preliminary and Critical Design Re-views (PDR and CDR) there is a general shift in the focus of mission design from aiding the vehicle design to providing mission specific products needed for pre-flight and real time operations. Some of the mission specific products need-ed include analysis of steering law performance, inputs into navigational accura-cy assessments, abort options at any point in the mission for each valid trajecto-ry in the launch window, recontact avoidance between the upper stage and Orion post nominal separation, etc.

Gutkowski, Jeffrey P.↗

NASA NTRS2019

Building Booster Separation Aerodynamic Databases for Artemis II

NASA’s Artemis II mission will mark the return of humans to near-lunar space for the first time since Apollo. Shortly after launch on the Space Launch System (SLS), a critical phase of ascent occurs when 16 small rockets fire to push the boosters away from the core. Minimizing the risk of failure during separation requires the construction of multiple 13-dimensional databases, including perturbations in position, flight conditions, and engine thrust. The SLS Computational Fluid Dynamics team used NASA’s FUN3D flow solver on the Pleiades and Electra supercomputers to run 5,780 simulations at nominal conditions and over 8,000 simulations with a core stage engine failure to generate the databases needed to verify the booster separation system for Artemis II.

Meeroff, Jamie G.↗