Vehicle and Mission Design of a Future Small Payload Launcher

The idea of space flight as an adventurous and exciting endeavor has been explored by popular culture since the early days of science fiction in the mid-1900s. These dreams can become true with establishment of an interplanetary travel facility that would support commercial space activities and research. Noticeably, an actual public demand for orbital trips, when they become commercially available, depends on overall price as well as specifics of the mission. This paper presents a design strategy for development of an interplanetary travel facility, with a goal to design a cost-effective and a fully autonomous Lunar Orbital Transportation (LOT) vehicle. The vehicle rendezvous between low earth orbit (LEO) and lunar orbit, capable of docking with a commercial space station, and can support a group of 9 people for a journey to the lunar orbit and back to the commercial space station. This paper also identifies/describes potential markets and related to them future missions in the era of space commercialization. The design will serve as a flight testbed and technology demonstration of critical elements and systems for deep space missions including human flight to Mars. The design of the LOT vehicle was based upon tourist’s itinerary and crew ConOps. Detailing the itinerary and the ConOps gave a better understanding of what crew and tourists are doing during these missions, and that further helped us to define functions and capabilities needed for a successful mission. Furthermore, design considerations for the required logistics and consumables specify area and volume requirements. The design of the vehicle will also require protection for specific space problems, such as micrometeorites, radiations, possible temperature variations, and space debris. After analyzing ConOps and tourism itineraries, a conclusion was made to have a minimum of three crew and a maximum of six tourists per LOT mission. This creates a 2:1 tourist-to-crew ratio, which is considered suitable for the crew to support tourists and ensure their safety.

PurposeThe purpose of this paper is to attempt an aerospaceplane design with the objective of Low-Earth-Orbit-and-Return-to-Earth (LEOARTE) under the constraints of safety, low cost, reliability, low maintenance, aircraft-like operation and environmental compatibility. Along the same lines, a “sister” point-to-point flight on Earth Suborbital Aerospaceplane is proposed.Design/methodology/approachThe LEOARTE aerospaceplane is based on a simple design, proven low risk technology, a small payload, an aerodynamic solution to re-entry heating, the high-speed phase of the outgoing flight taking place outside the atmosphere, a propulsion system comprising turbojet and rocket engines, an Air Collection and Enrichment System (ACES) and an appropriate mission profile.FindingsIt was found that a LEOARTE aerospaceplane design subject to the specified constraints with a cost as low as 950 United States Dollars (US$) per kilogram into Low Earth Orbit (LEO) might be feasible. As indicated by a case study, a LEOARTE aerospaceplane could lead, among other activities in space, to economically viable Space-Based Solar Power (SBSP). Its “sister” Suborbital aerospaceplane design could provide high-speed, point-to-point flights on the Earth.Practical implicationsThe proposed LEOARTE aerospaceplane design renders space exploitation affordable and is much safer than ever before.Originality/valueThis paper provides an alternative approach to aerospaceplane design as a result of a new aerodynamically oriented Thermal Protection System (TPS) and a, perhaps, improved ACES. This approach might initiate widespread exploitation of space and offer a solution to the high-speed “air” transportation issue.

VentureStartm… reaping the benefits of the X-33 program

We describe the use of the Cyber-Physical Modeling Language (CyPhyML) to support trade studies and integration activities in system-level vehicle designs. CyPhyML captures integration interfaces across multiple design domains for system components, and generic design assembly rules given in terms of architecture alternatives. The CyPhyML tools support automated exploration of system-level architectural and parametric tradeoffs using a suite of design exploration tools that can be applied to models at different levels of fidelity and scale. Our overall approach includes exploration over the space of potential designs by evaluating structural combinations and then comparison of designs by simulating the dynamics of systems and subsystems with varying degrees of detail. In that flow, we use the DESERT toolkit for design tradeoffs that can be evaluated from the structure of the design model (i.e. interconnections of components and their parameters). DESERT extends a graphical modeling language with concepts and relations to define structural alternatives and constraints on system properties. Alternatives are given in an abstract way, decoupled from the details of the encoding of the combinatorial design space problem for the underlying binary decision diagram (BDD) solver. For desirable design instance models, the tools automatically assemble complete vehicle or subsystem computer-aided design (CAD) models from the associated component CAD models, and likewise create detailed finite-element structural analysis (FEA) models for selected component assemblies. Evaluation results are presented in a dashboard display which provides comparisons between different valid designs over all specified design metrics.

The possibility to launch small payloads into LEO by means of an electromagnetic railgun is the motivation of the present study. Here it is assumed that an electromagnetic railgun with barrel length of 180 m is able to launch a 120 mm diameter projectile at a speed of 6 km/s. In order to the payload achieves orbital motion and to compensate the velocity losses due to gravity and aerodynamic drag, a rocket engine which supplies the required velocity difference is integrated into the projectile structure. The development of such propelled projectile requires key technologies that could also benefit other industrial sectors. The paper presents the main concept and design aspects to be considered for such propelled projectile with special emphasis on: flight mechanics, aero/thermodynamics, materials and propulsion characteristics. The aerodynamic form of the projectile is optimized to achieve minimum drag and lower heat loads. Launch angles and trajectory optimization analysis are carried out by means of 3 degree of freedom simulations (3DOF). Furthermore, three concurrent propulsions systems are here considered for the rocket engine: solid, liquid and hybrid propellants type. The result of analysis pointed that the most promising propulsion systems are: (I) hybrid engine based on hydrogen peroxide/wax propellants and (II) liquid engine based on hydrogen peroxide/hydrocarbon propellants. The structural mass is analyzed mostly on the basis of carbon fiber reinforced materials as well as classical aerospace metallic in small amount. Finally, the paper gives a critical mark for the technical feasibility of small rockets for such missions.

The National Aeronautics and Space Administration (NASA) is preparing to send humans beyond Low Earth Orbit and eventually to the surface of Mars. As part of the Evolvable Mars Campaign, different vehicle configurations are being designed and considered for delivering large payloads to the surface of Mars. Weight and packing volume are driving factors in the vehicle design, and the thermal protection system (TPS) for planetary entry is a technology area which can offer potential weight and volume savings. The feasibility and potential benefits of a ceramic matrix composite hot structure concept for different vehicle configurations are explored in this paper, including the nose cap for a Hypersonic Inflatable Aerodynamic Decelerator (HIAD) and an aeroshell for a mid lift-to-drag (Mid L/D) concept. The TPS of a planetary entry vehicle is a critical component required to survive the severe aerodynamic heating environment during atmospheric en- try. The current state-of-the-art is an ablative material to protect the vehicle from the heat load. The ablator is bonded to an underlying structure, which carries the mechanical loads associated with entry. The alternative hot structure design utilizes an advanced carbon-carbon material system on the outer surface of the vehicle, which is exposed to the severe heating and acts as a load carrying structure. The preliminary design using the hot structure concept and the ablative concept is determined for the spherical nose cap of the HIAD entry vehicle and the aeroshell of the Mid L/D entry vehicle. The results of the study indicate that the use of hot structures for both vehicle concepts leads to a feasible design with potential weight and volume savings benefits over current state-of-the-art TPS technology that could enable future missions.

The ubiquitous nature of microbiology is reflected in the diversity of microbiological research and operational efforts at NASA. For example, the impact of microorganisms on other planets and the protection of Earth from the potential of microbial life elsewhere is the responsibility of the Office of Planetary Protection (http://planetaryprotection.nasa.gov/ about). While the Office of Planetary Protection does not include forward or back contamination to or from a low Earth orbit, research platforms, such as the ISS, are being used to better understand the survival of microorganisms and corresponding contamination control in the extreme conditions of space (2, 4, 10). Another example is the NASA Center for Astrobiology, which focuses on the origin, evolution, distribution, and future of life in the universe (https:// astrobiology.nasa.gov/). The large focus on microbiology is within its human exploration operations. NASA has historically set microbiological requirements, including stringent monitoring regimes, to mitigate risks to the health and performance of astronauts. Microorganisms can have both positive and negative impacts on many aspects of human spaceflight, including the risk and prevention of infectious diseases, performance of Environmental Control and Life Support Systems (ECLSS), spaceflight foods, and vehicle design and integrity. Even though a great amount of information has been obtained, several key questions regarding the impact of microorganisms on human spaceflight still remain. Research into the uncertainties of risks that may affect crew health are the responsibility of the NASA Human Research Program (http://www.nasa.gov/exploration/human research/). The research that addresses our fundamental understanding of space life science and its translation for benefits to the general public on Earth is the responsibility of NASA Space Biology.

Index allocation for a reusable LOX/CH4 rocket engine

The Ballistic Missile Defense System (BMDS) is a layered system incorporating elements in space. In addition to missile warning systems at geosynchronous altitudes, an operational BMDS will include a low Earth orbit (LEO) system-the Space Tracking and Surveillance System (STSS). It will use infrared sensing technologies synergistically with the Space Based Infrared Systems (SBIRS) and will provide a seamless adjunct to radars and sensors on the ground and in airborne platforms. STSS is being designed for a future operational capability to defend against evolving threats. STSS development is divided into phases, commencing with a two-satellite demonstration constellation scheduled for launch in 2008. The demonstration satellites will conduct a menu of tests and experiments to prove the system concept, including the ground segment. They will have limited operational capability within the integrated BMDS. Data from the demonstration satellites will be received and processed by the Missile Defense Space Experiment Center (MDSEC), a part of the Missile Defense Integration and Operations Center (MDIOC). MDA launched in 2007 into LEO a satellite (NFIRE) designed to make near-field multispectral measurements of boosting targets and to demonstrate laser communication, the latter in conjunction with the German satellite TerraSAR-X. The gimbaled, lightweight laser terminal has demonstrated on orbit a 5.5 gbps rate in both directions. The filter passbands of NFIRE are similar to the STSS demonstrator track sensor. While providing useful phenomenology during its time on orbit, NFIRE will also serve as a pathfinder in the development of STSS operations procedures.

Multi-disciplinary optimization of a space re-entry vehicle using skeleton-based integral soft objects

Pre-engineering spaceflight validation of environmental models and the 2005 HZETRN simulation code

A simplified trajectory analysis model for small satellite payload recovery from low Earth orbit

Our goal in this work is to develop analytical tools to support the definition of balanced and compatible sets of vehicle specifications in the early stages of vehicle development. In this paper, we discuss the development and application of a Technical Feasibility Model (TFM) that may be used in preliminary design to assess the technical feasibility and optimality of specified combinations of vehicle performance targets. For this paper, we have exercised the TFM specifically to explore the relationships between vehicle mass, vehicle performance measures, (such as acceleration, fuel efficiency, and interior roominess), and high-level vehicle design parameters (such as overall exterior dimensions, occupant positions, and selection of a powertrain). The TFM is developed by first applying a MultiObjective Genetic Algorithm to a multidisciplinary design framework to generate a set of Pareto-optimal design solutions, then applying response surface methods to generate a smooth mathematical representation of the Pareto set, and finally using geometric construction to analyze the position of a test point relative to the representation of the Pareto set. Results of this analysis include an assessment of the feasibility and optimality of the test point as well as a variety of projections from the test point to the representation of the Pareto set that may be used to identify opportunities for refining, relaxing, improving, or prioritizing performance specifications. The mapping between performance space and design space has been preserved, allowing for investigation of relationships between performance specifications and design variable settings.

This paper proposes an approach for the multidisciplinary design & optimization of space launch vehicle, wherein the structure/mass, aerodynamics, propulsion and the trajectory system design are performed simultaneously using non-dominated sorting based genetic algorithm. The design problem is posed as an optimization problem in which all the discipline parameters constitute the design variables, which are used to optimize the multi-objective function, thereby resulting in an optimal overall design. The approach is demonstrated by application to the integrated design of a four stage solid launch vehicle. The mission is to deliver predefined payload to the low earth (circular) orbit. An existing, real world system has been used to validate the multistage launch vehicle system model.

Thanks to recent considerable progress in proton exchange membrane (PEM) technology, fuel cells and electrolyzers are on the verge of widespread commercialization. When a fuel cell and an electrolyzer are combined, a regenerative fuel cell (RFC) system is formed. By using an auxiliary power supply, such as solar power, for recharging, an RFC provides a complete power system for niche markets such as low-earth-orbit (LEO) satellites. The thermodynamics of RFC systems are presented, and design tradeoffs are investigated: a unitized system, where the fuel cell and the electrolyzer are combined into a single electrochemical device, is compared to a discrete system, where the fuel cell and the electrolyzer are discrete components. The analyses show that the RFC is well suited for LEO space applications, due to an appropriate charge/discharge cycle, and represents a high-energy-density, high-efficiency power solution.