Space Launch System Development Status and Advanced Capability for Exploration Missions

NASA's Marshall Space Flight Center (MSFC) is directing efforts to build the Space Launch System (SLS), a heavy-lift rocket that will carry the Orion Multi-Purpose Crew Vehicle (MPCV) and other important payloads far beyond Earth orbit (BEO). Its evolvable architecture will allow NASA to begin with Moon fly-bys and then go on to transport humans or robots to distant places such as asteroids and Mars. Designed to simplify spacecraft complexity, the SLS rocket will provide improved mass margins and radiation mitigation, and reduced mission durations. These capabilities offer attractive advantages for ambitious missions such as a Mars sample return, by reducing infrastructure requirements, cost, and schedule. For example, if an evolved expendable launch vehicle (EELV) were used for a proposed mission to investigate the Saturn system, a complicated trajectory would be required - with several gravity-assist planetary fly-bys - to achieve the necessary outbound velocity. The SLS rocket, using significantly higher characteristic energy (C3) energies, can more quickly and effectively take the mission directly to its destination, reducing trip time and cost. As this paper will report, the SLS rocket will launch payloads of unprecedented mass and volume, such as “monolithic” telescopes and in-space infrastructure. Thanks to its ability to co-manifest large payloads, it also can accomplish complex missions in fewer launches. Future analyses will include reviews of alternate mission concepts and detailed evaluations of SLS figures of merit, helping the new rocket revolutionize science mission planning and design for years to come.

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 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.

Science on the lunar surface facilitated by low latency telerobotics from a Lunar Orbital Platform - Gateway

Aerospace - Rocket. The Gallery. Nasa's Space Launch System

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 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.

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.

The ram accelerator, a chemically driven ramjet-in-tube device is a new option for direct launch of acceleration-insensitive payloads into earth orbit. The projectile is the centerbody of a ramjet and travels through a tube filled with a premixed fuel-oxidizer mixture. The tube acts as the cowl of the ramjet. A number of new concepts for a ram accelerator space launch system are presented. The velocity and acceleration capabilities of a number of ram accelerator drive modes, including several new modes, are given. Passive (fin) stabilization during atmospheric transit is investigated and found to be promising. Gasdynamic heating in-tube and during atmospheric transit is studied; the former is found to be severe, but may be alleviated by the selection of the most suitable drive modes, transpiration cooling, or a hydrogen gas core in the launch tube. To place the payload in earth orbit, scenarios using one impulse and three impulses (with an aeropass) and a new scenario involving an auxiliary vehicle are studied. The auxiliary vehicle scenario is found to be competitive regarding payload, and requires a much simpler projectile, but has the disadvantage of requiring the auxiliary vehicle.

The problem of trajectory optimization is important in all space missions. The solution of this problem enables one to specify the optimum thrust steering program which should be followed to achieve a specified mission objective, simultaneously satisfying the constraints.1 It is well known that whether or not the ascent trajectory is optimal can have a significant impact on propellant usage for a given payload, or on payload weight for the same gross vehicle weight.2 Consequently, ascent guidance commands are usually optimized in some fashion. Multi-stage vehicles add complexity to this analysis process as changes in vehicle properties in one stage propagate to the other stages through gear ratios and changes in the optimal trajectory. These effects can cause an increase in analysis time as more variables are added and convergence of the optimizer to system closure requires more analysis iterations. In this paper, an approach to simplifying this multi-stage problem through the creation of an upper stage capability boundary is presented. This work was completed as part of a larger study focused on trade space exploration for the advanced booster system that will eventually form a part of NASA s new Space Launch System.3 The approach developed leverages Design of Experiments and Surrogate Modeling4 techniques to create a predictive model of the SLS upper stage performance. The design of the SLS core stages is considered fixed for the purposes of this study, which results in trajectory parameters such as staging conditions being the only variables relevant to the upper stage. Through the creation of a surrogate model, which takes staging conditions as inputs and predicts the payload mass delivered by the SLS upper stage to a reference orbit as the response, it is possible to identify a of staging conditions which all satisfy the SLS requirement of placing 130 metric tons into low-Earth orbit (LEO).3 This identified surface represents the 130 metric ton capability boundary for the upper stage, such that if the combined first stage and boosters can achieve any one staging point on that surface, then the design is identified as feasible. With the surrogate model created, design and analysis of advanced booster concepts is streamlined, as optimization of the upper stage trajectory is no longer required in every design loop.

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.

Capabilities for a 2033 Crew Mars Flyby Mission Launched with the NASA Space Launch System

As long-distance human space travel becomes more realistic, the need for long-term storage of cryogenic propellants becomes a forefront issue. Effective cryo-fluid system management becomes a necessity to solve this issue. The fuel–oxidizer combination of choice for NASA’s Space Launch System is liquid hydrogen () and liquid oxygen (LOX). The current strategy used on many upper stage vehicles, including United Launch Alliance’s Advanced Cryogenic Evolved Stage, is to use the boil-off gases of the colder cryogen () to help insulate the warmer cryogen (LOX) and reduce the static boil-off. In this work, experimental measurements are collected for a range of catalyst materials and weight loadings of and catalysts on Nomex scrim blankets. The blankets are configured in a nonisothermal catalytic reactor with a 23 K input temperature and heated to 90 K output temperature. Results of catalyzation of parahydrogen to orthohydrogen gas are presented as well as a catalyzation model to represent the measurements for future use. In addition, simplified computational fluid dynamics studies have been initiated to estimate the surface-based conversion rate constant by matching numerical results and experimental data.

This paper describes the successful development and test of a multipoint fiber optic hydrogen microsensors system during the static firing of an Evolved Expandable Launch Vehicle (EELV)/Delta's common booster core (CBC) rocket engine at NASA's Stennis Space Center. The hydrogen sensitive chemistry is fully reversible and has demonstrated a response to hydrogen gas in the range of 0% to 10% with a resolution of 0.1% and a response time of ≤5 seconds measured at a gas flow rate of 1 cc/min. The system consisted of a reversible chemical interaction causing a change in reflective of a thin film of coated Palladium. The sensor using a passive element consisting of chemically reactive microcoatings deposited on the surface of a glass microlens, which is then bonded to an optical fiber. The system uses a multiplexing technique with a fiber optic driver-receiver consisting of a modulated LED source that is launched into the sensor, and photodiode detector that synchronously measures the reflected signal. The system incorporates a microprocessor to perform the data analysis and storage, as well as trending and set alarm function. The paper illustrates the sensor design and performance data under field deployment conditions.