Global Comparison of CFD and Wind Tunnel Derived F&M Databases for the Space Launch System

Until 1998, the federal government has been the predominant purchaser of space launches in the U.S. through the purchase of hardware and services. Historically, the government provided the necessary infrastructure for launches from the federal DoD and NASA launch ranges. In this historical model, the federal government had complete ownership, responsibility, liability, and expense for launch activities. In 1998, commercial space launches accounted for 60% of U.S. launches. This growth in commercial launches has increased the demand for launch range services. However, the expense, complexity of activities, and issues over certification of flight safety have deterred the establishment of purely commercial launch sites, with purely commercial being defined as without benefit of capabilities provided by the federal government. Provisions of the Commercial Space Launch Act have enabled DoD and NASA to support commercial launches from government launch ranges on a cost-reimbursable, non-interference basis. The government provides services including use of facilities, tracking and data services, and range and flight safety. In the 1990’s, commercial space market projections indicated strong potential for large numbers of commercial satellites to be launched well into the first decade of the 21st century. In response to this significant opportunity for economic growth, several states established spaceports to provide the services necessary to meet these forecast commercial needs. In 1997, NASA agreed to the establishment of the Virginia Space Flight Center (VSFC), a commercial spaceport, at its Wallops Flight Facility. Under this arrangement, NASA agreed to allow the Virginia Commercial Space Flight Authority (VCSFA) to construct facilities on NASA property and agreed to provide launch range and other services in accordance with the Space Act and Commercial Space Launch Act in support of VSFC launch customers. A partnership relationship between NASA and VCSFA has emerged which pairs the strengths of the established NASA Test Range and the state-sponsored, commercial launch facility provider in an attempt to satisfy the needs for flexible, low-cost access to space. The continued viability of the VSFC and other commercial spaceports depend upon access to a space launch and re-entry range safety system that assures the public safety and is accepted by the public and government as authoritative and reliable. DoD and NASA budget problems have resulted in deteriorating services and reliability at federal ranges and has caused fear with respect to their ability to service the growing commercial market. Numerous high level studies have been conducted or are in progress that illuminate the deficiencies. No federal agency has been provided the necessary funding or authority to address the nations diminishing space launch capability. It is questionable as to whether the U.S. can continue to compete in the global space launch market unless these domestic space access problems are rapidly corrected. This paper discusses a potential solution to the lack of a coordinated response in the U.S. to the challenge presented by the global market for space launch facilities and services.

Article analyzes the changes in the space economy, as well as the U.S. federal policy for preserving American domination in this area. The evolution of the space economy in the United States is briefly considered, including formation of venture businesses since late 1990s (due to the computer and Internet revolution and development of space technologies). Of particular importance is a so-called Astropreneurship – growth of startups, developing space launch systems and spacecrafts. Amid technological factors, the main driver for the rise of Astropreneurship was a suboptimal technological and market strategy of Boeing and Lockheed Martin duopoly. It created conditions for disruptive innovation development in the industry and transition to a new, more open model of innovation processes there. Although in the 1980–1990’s federal government enacted a business-friendly regulations for the commercial space sector and set measures to support it through public procurement, a new situation in the markets and in sectoral innovation system has become a challenge for national policy, enforced by budget constraints and other factors. Catastrophe of the Columbia space shuttle in 2003, which raised the issue of maintaining independent U.S. access to the International Space Station and, in the future, to other orbital operations after shuttles’ phase-out, triggered changes in federal policies for the industry. As a result, since 2006 NASA (with some DoD support) initiated several programs to develop new space launch and delivery systems. These programs were based on the public-private partnerships with active involvement of small and medium innovative enterprises, primarily startups. The results turned out to be more significant than originally intended: formation of a new model of federal industrial and technological policy in the space sector, almost similar to the DARPA principles. The new policy assumes the role of NASA and, more broadly, the federal government as an equal (rather than a dominant) participant in space industry innovative processes and as a more active organizer of the industry`s innovation system (considering changes in its nature). Despite all activities are mission-oriented, more attention is paid to the development of complex partnerships, ecosystems, etc. – with an emphasis both on satisfaction of government needs and on ensuring market leadership of the U.S. companies. However, this new model faces sever al problems: ambiguity of federal expenditures on space- related research and development; need to restructure NASA science and technology organization; efficiency of federal efforts in support of new partners and processes – taking into account specifics of available policy instruments. The future of federal efforts will be determined by the solution of these problems and by the need to extend new model and practices to the key (in terms of market size and technology advances) segments of space economy – satellite manufacturing and the use of space data.

The National Aeronautics and Space Administration (NASA) is developing new capabilities for human and scientific exploration beyond Earth's orbit. This effort includes the Space Shuttle derived Space Launch System (SLS), the Multi-Purpose Crew Vehicle (MPCV) and the Ground Systems Development and Operations (GSDO). There are several requirements and Technical Performance Measures (TPMs) that have been levied by the Exploration Systems Development (ESD) upon the SLS, MPCV, and GSDO Programs including an integrated Launch Availability (LA) TPM. The LA TPM is used to drive into the SLS, Orion and GSDO designs a high confidence of successfully launching exploration missions that have narrow Earth departure windows. The LA TPM takes into consideration the reliability of the overall system (SLS, Orion and GSDO), natural environments, likelihood of a failure, and the time required to recover from an anomaly. A challenge with the LA TPM is the interrelationships between SLS, Orion, GSDO and the natural environments during launch countdown and launch delays that makes it impossible to develop an analytical solution for calculating the integrated launch probability. This paper provides an overview of how Discrete Event Simulation (DES) modeling was used to develop the LA TPM, how it was allocated down to the individual programs, and how the LA analysis is being used to inform and drive the SLS, Orion, and GSDO designs to ensure adequate launch availability for future human exploration.

Space launch projects are extremely risky, and any equipment failure or human error may lead to disastrous consequences. Probabilistic risk assessment (PRA) is beneficial to qualitative analysis of risk, but it has not been paid enough attention in risk analysis for space launch systems (SLSs). Compared with most qualitative risk analysis in this field, this paper proposes a risk analysis framework based on Bayesian network (BN) with fuzzy method, which is suitable for probabilistic risk analysis of SLS. This method establishes a risk analysis model of SLS based on statistics and expert experience and reduces the uncertainty of the model by using fuzzy theory. By predicting the system risk probabilities, diagnosing the key risk causes, determining the risk conduction path, and performing a sensitivity analysis, the proposed risk analysis framework is aimed at alleviating this drawback to deal more effectively with the uncertainties in the field of space launches. A case study of space launches demonstrates and verifies the proposed method, and it also provides guidance for similar engineering projects.

Scientific exploration of the solar system has traditionally been the realm of large and expensive projects known as primary missions. These missions are usually costly, and they tend to have a long development cycle of massive spacecraft carrying scientific payloads. The maturation of mini- and micro-satellites over the past 40 years has allowed space science to be carried out in lower cost missions. In more recent years, the space community has started to consider using even nanosatellite platforms (such as CubeSats that are typically within a few tens of kilograms) for scientific endeavor in the solar system. Robotic exploration of the solar system using nanosatellites is attractive because of their lower cost, less stringent risk posture, and a shorter development schedule compared to traditional primary science missions. Therefore, it is beneficial to facilitate the development of nanosatellite technologies for deep space exploration by providing an opportunity for a largescale demonstration. Motivated by this, NASA is offering free space launch services on Space Launch System (SLS) to 13 deep space 6U CubeSats, as well as is providing the ground network support if the selected CubeSat missions ask for it. The CubeSats have been chosen competitively as secondary payloads on the SLS launch vehicle carrying the Exploration Mission-1 (EM-1), which will be the first integrated flight test of NASA's Deep Space Exploration Systems including the Orion spacecraft, SLS rocket, and the newly upgraded Exploration Ground Systems at Kennedy Space Center in Cape Canaveral, FL, USA. The 13 EM-1 CubeSats are currently going through the final stages of integration and testing to be ready for launch in late 2020. To help showcase the potential impact of the EM-1 CubeSats, this Special Issue within IEEE AEROSPACE AND ELECTRONIC SYSTEMS MAGAZINE (AESM) is organized to present their science objectives, development story, the challenges faced, and lessons learned by the development teams to share with the community. The three Editors of the Special Issue (SI) would like to thank all the authors from the EM-1 CubeSat teams who managed to make our publication schedule, as well as authors from the EM-1 primary mission team and DSN team. The Editors would also like to thank AESM for offering the publication opportunities to this SI and allowing our SI to come in three parts due to a large volume of papers we have to include. The Part 1 of the SI consists of five papers that are briefly introduced.

The National Aeronautics and Space Administration s (NASA) Space Launch System (SLS) Program, managed at the Marshall Space Flight Center (MSFC), is charged with delivering a new capability for human and scientific exploration beyond Earth orbit (BEO). The SLS may also provide backup crew and cargo services to the International Space Station, where astronauts have been training for long-duration voyages to destinations such as asteroids and Mars. For context, the SLS will be larger than the Saturn V, providing 10 percent more thrust at liftoff in its initial 70 metric ton (t) configuration and 20 percent more in its evolved 130-t configuration. The SLS Program knows that affordability is the key to sustainability. This paper will provide an overview of its operations strategy, which includes initiatives to reduce both development and fixed costs by using existing hardware and infrastructure assets to meet a first launch by 2017 within the projected budget. It also has a long-range plan to keep the budget flat using competitively selected advanced technologies that offer appropriate return on investment. To arrive at the launch vehicle concept, the SLS Program conducted internal engineering and business studies that have been externally validated by industry and reviewed by independent assessment panels. A series of design reference missions has informed the SLS operations concept, including launching the Orion Multi-Purpose Crew Vehicle (MPCV) on an autonomous demonstration mission in a lunar flyby scenario in 2017, and the first flight of a crew on Orion for a lunar flyby in 2021. Additional concepts address the processing of very large payloads, using a series of modular fairings and adapters to flexibly configure the rocket for the mission. This paper will describe how the SLS, Orion, and Ground Systems Development and Operations (GSDO) programs are working together to create streamlined, affordable operations for sustainable exploration for decades to come.

As NASA focuses on Mars mission readiness, a crewed lunar lander concept may be timely, as there are both Mars exploration and lunar exploration motives for its development. A lunar lander can be used as a Mars lander technology demonstrator, including demonstrating technologies used in the final phase of Mars descent/landing, as well as technologies/techniques for habitation and ascent to orbit. Utilizing the cislunar Deep Space Gateway and Space Launch System (SLS) transportation capabilities, the addition of a lander would enable international and scientific communities to explore the lunar surface The Boeing lunar lander concept supports a crew of four on a two week (one lunar day) surface sortie and is extensible to longer duration missions with the addition of separately delivered surface assets such as a mobile habitat. The lander is optimized to take advantage of the cislunar Deep Space Gateway architecture; the lander elements are delivered using the SLS Block 2 cargo launch capability and the SLS Orion/cargo co-manifest capability. Overall mass of the ascent module was minimized by allocating habitation and crew surface access functions to the descent module. Recurring costs were reduced by designing a reusable ascent module. This paper will describe the lander design and operation concept, key trade studies and rationale for the selected options. Trade studies were conducted in areas such as spacecraft elements, engines, propellant, airlock, and life support subsystems. Analysis of the Deep Space Gateway orbit is discussed with respect to its compatibility with: a) surface landing operations, b) earth-to-orbit element/cargo delivery operations, and c) access to various lunar surface regions. The paper will show that use of the SLS, Orion and the Deep Space Gateway in support of a human lunar lander exploration architecture is feasible and practical and will explain how the design described can be extensible to future Mars lander designs.

The NASA Space Launch System (SLS) Block 1 configuration is progressing toward a 2021 first flight, which will launch an uncrewed Orion capsule to the Moon. The SLS' Core Stage has finished its Green Run, full duration test at NASA Stennis Space Center. In this report, the SLS development status and its launch capabilities are discussed, and the new optimized Exploration Upper Stage, currently in development, which will optimize the SLS Core and Booster Stages, is presented. Ten Beyond Earth Orbit (BEO) missions, enabled or enhanced by the SLS, are also reviewed.

Exciting and robust human and robotic exploration missions will be enabled by the NASA Space Launch System (SLS). The SLS's capabilities for launching heavy payloads with large diameters (up to 10 meters) will enable a variety of science rich missions including Mars and outer planet missions; including Human missions utilizing the Orion Crew Vehicle. Missions to cislunar destinations will serve as proving grounds for later Crewed Mars Missions in the 2030s. In this report an outline is given of the capabilities of the SLS, strategies for exploration missions are presented, and descriptions of SLS based payload elements are given. The SLS program is making progress toward delivering a new capability for Beyond Earth Orbit exploration.

The Boeing Company is designing, developing and manufacturing the Core Stage and Avionics for NASA's heavy lift rocket, the Space Launch System (SLS), which will support missions beyond low earth orbit for the first time since the Apollo days. On the SLS vehicle many sensors are selected and used to provide critical information for testing, ground operations, and vehicle flight environments. Appropriate data acquisition systems that can interface and process the selected sensors is important to the success of the program. In this program, multiple data acquisition systems are used to interface with sensors at various development stages of the SLS vehicle, and at different sensor criticality levels. These phased configurations of data acquisition classes and complexity range from gathering sensor data during system level testing at Stennis Space Center (SSC), to transmitting flight information after liftoff at the Kennedy Space Center. One challenge that engineers are faced with pertains to how the data acquisition system is defined prior to understanding the quantity and type of sensors required for the application. This paper focuses mainly on the SLS flight system and provides a general overview of the system architecture. The paper also addresses the strategy used to define the data acquisition system, along with the methodology for selecting data acquisition cards that fit with the application.

NASA’s Space Launch System (SLS) and Orion Crew Module have been studied in wind tunnels using the planar laser-induced fluorescence (PLIF) technique for flow visualization and quantitative thermometry and in flight using high-resolution imagery and quantitative infrared surface heating measurements This paper reviews both types of measurements.

A wind-tunnel investigation of a 0.009 scale model of the Space Launch System (SLS) was conducted in the NASA Langley Unitary Plan Wind Tunnel to characterize the aerodynamics of the core and solid rocket boosters (SRBs) during booster separation. High-pressure air was used to simulate plumes from the booster separation motors (BSMs) located on the nose and aft skirt of the SRBs. Force and moment data were acquired on the core and SRBs. These data were used to corroborate computational fluid dynamics (CFD) calculations that were used in developing a booster separation database. The SRBs could be remotely positioned in the x-, y-, and z-direction relative to the core. Data were acquired continuously while the SRBs were moved in the axial direction. The primary parameters varied during the test were: core pitch angle; SRB pitch and yaw angles; SRB nose x-, y-, and z-position relative to the core; and BSM plenum pressure. The test was conducted at a free-stream Mach number of 4.25 and a unit Reynolds number of 1.5 million per foot.

An electromagnetic launching system presents a viable projectile propulsion alternative with low cost and minimal environmental drawbacks. A coil gun system propels a projectile using an electromagnetic force and the system is mainly employed in military weapon systems and space launch systems. In this paper, we perform optimization design to improve the muzzle velocity by analyzing the sensitivity. The muzzle velocity, which is the most important design function variable, is affected by design variables including the number of axial turns in the electromagnetic coil, number of radial turns in the electromagnetic coil, initial distance between the projectile and the coil, inner radius of the electromagnetic coil, and length of the projectile. An orthogonal arrays matrix is configured, and a finite element analysis is performed utilizing the commercial electromagnetic analysis software MAXWELL. The muzzle velocity of the optimal design is 62.4% greater than that of the initial design.

The Johns Hopkins University Applied Physics Laboratory (APL) has been tasked by the NASA Heliophysics Division to (re-)study a robotic Interstellar Probe mission. The top-level requirement is to provide input to support the next Solar and Space Physics “Decadal Survey” in the United States, with a nominal time frame of performance from 2023 to 2032.The study approach is to look widely across both relevant scientific and technical communities and assemble a “Menu” of what has been, and what can be, done with respect to Interstellar Probe desires and concepts past. By its nature this assemblage is a “superset” of what might be implemented; “ordering” from the menu will be a charge to a future Science Definition Team – at NASA’s discretion. This approach has been adopted successfully in the past, with the emphasis on informing the Survey participants of valid possibilities, while not dictating a “best” solution.The components of such a menu are not random. They must flow from compelling top-level science goals, through explicit measurement requirements, instrumentation to make those measurements, and an assessment of how those measurements have provided “closure” to the investigation, i.e. have addressed the science goals.For this type of engineering “menu” study, payload instrument possibilities and capabilities, with representative masses and power requirements, are required to help assess the overall spacecraft cost, mass, and achievable speed for a given launch vehicle configuration. Details have and will continue to evolve as the engineering aspects mature and as more people throughout the international scientific community continue to contribute their ideas.Neither critical trade-offs nor enabling technologies are new. For example, with a given launch system, the total energy that can be imparted to the spacecraft is fixed. Thus, with a given set of planetary gravity assists there is a trade between the maximum asymptotic escape speed from the Sun and the total mass of the spacecraft, which, in turn, tends to scale with the payload mass. As a starting point, a range of spacecraft masses from ~300 to 800 kg, corresponding roughly to Pioneer (251.8 kg) through Voyager (825.4 kg), has been considered. Communication has focused on microwave downlink (X-band or Ka-band), which is well developed, known, and robust. While optical laser comm might achieve far higher downlink rates, it requires extreme pointing stability, and the associated lifetime also needs continuing investigation. The need and capability for radioisotope power systems (RPS) for powering deep-space robotic spacecraft operating far from the Sun is well established  and will be required here as well. Also, in this study the use of the Space Launch System (SLS) cargo version is in use along with upper stages. Other “lunar capable” launch vehicles have also been “spot checked,” but more complete performance data is needed from their respective vendors for a definitive assessment.As with any other study of this type, initial engineering requirements must be imposed to begin an inherently iterative process of design. Engineering requirements are needed to frame the engineering study and “bound the box” – but allow for trades. As with other aspects of this study, these are also still evolving:(1)       Enable a mission that can be launched no later than 1 January 2030 (technology driver)(2)       Have the capability to operate from a maximum range of not less than (NLT) 1000 astronomical units (a.u.) from the Sun (communications system driver)(3)       Require no more than 600 Watts of electrical power (We) at the beginning of mission (BOM) and be able to operate at no less than half of the BOM amount at the end of mission (EOM) (power system driver)(4)       Achieve a mission lifetime of NLT 50 years with an “acceptable” probability of success (drives parts program, physics of failure analyses, programmatics, and policy)Given the requirement of near-term flight, we focus entirely upon ballistic solutions with Jupiter Gravity Assists (JGA) in order to maximize the asymptotic solar system escape speed. Three options are under study: The first two use prograde gravity assists at Jupiter, one passive and one active with an upper stage burn in Jupiter’s gravity well; the third uses a retrograde gravity assist and a powered “Oberth maneuver” near the Sun.The current focus is upon a medium-mass, all-heliophysics payload to obtain a first cut of basic architecture approach and initial subsystem master equipment list (MEL) with power requirements and margins and reserves, as appropriate. This is being used to size the communication system, avionics, and guidance and control and data systems. It will also serve as a baseline for looking at both upscale and downscale payload and capability options, and trades with planetary and astrophysics science options, including effects on required spacecraft concept of operations and the payload mass, power, and cost tradespaces.The current effort is well into this initial concept definition via seven focus studies:1) Longevity - Spacecraft lifetimes/failures, long-lasting systems, failure modes, ground operations, staff renewal, and knowledge documentation and retention 2) Instruments - Candidate payload components with parameters + operating requirements 3) Trajectory and launch vehicle trades including Jupiter’s radiation and dust ring keep-out zone(s) 4) Communication and guidance and control trades, achievable pointing accuracy 5) Thermal shield requirements and possibilities for a near-Sun maneuver, including quantifying shield mass, radiation pressure effects, and center-of-mass migration and attitude control requirements during a near-Sun, rocket-motor burn 6) Overall mechanical layout, fairing clearances, payload adaptors, instrument clear fields of view, boom and antenna options, deployments, and associated pointing control 7) Power requirements, including payload and spacecraft systems, parasitic heating of propellant lines, surge capability trades, and shunt requirements for use with a Next Generation (NG)-RPS The study continues to solicit input on all of these topics from the space science and engineering communities via meetings and workshops. We remain on track for delivering the final report to NASA’s Heliophysics Division in late calendar year 2021.

A human mission to Mars has long been an exploration goal of the United States and the international community. To achieve this goal, NASA is developing the Orion crew capsule and Space Launch System (SLS) as key elements in the eventual architecture for the Evolvable Mars Campaign (EMC). Both elements are designed to advance human spaceflight beyond Low Earth Orbit (LEO) for the first time since the Apollo missions with the eventual goal of landing humans on Mars. This aggressive goal is complicated by the need to optimize the use of government resources during a period when budgets are challenged. Defining and prioritizing the technology and hardware needed is crucial to achieving the goal of landing humans on Mars. This paper will discuss potential near term missions that leverage Orion and SLS. A unique feature of the SLS/Orion system is supporting co-manifested payloads along with launching crews beyond LEO. Doing so provides new flexibility in achieving human exploration, operations, and science objectives on any given mission. Utilizing this co-manifest capability, a number of potential near term missions have been identified that will mature the elements needed for an eventual Mars mission, develop long duration mission operations techniques necessary for the Mars mission, refine life support systems, and conduct science investigations that were previously impractical. This paper will discuss a possible roadmap of missions that could be conducted over the next decade to place humans on the path to Mars.