Application of Risk within Net Present Value Calculations for Government Projects

In January 2004, President Bush announced a new vision for space exploration. This included retirement of the current Space Shuttle fleet by 2010 and the development of new set of launch vehicles. The President's vision did not include significant increases in the NASA budget, so these development programs need to be cost conscious. Current trade study procedures address factors such as performance, reliability, safety, manufacturing, maintainability, operations, and costs. It would be desirable, however, to have increased insight into the cost factors behind each of the proposed system architectures. This paper reports on a set of component trade studies completed on the upper stage engine for the new launch vehicles. Increased insight into architecture costs was developed by including a Net Present Value (NPV) method and applying a set of associated risks to the base parametric cost data. The use of the NPV method along with the risks was found to add fidelity to the trade study and provide additional information to support the selection of a more robust design architecture

NASA Experience with Pogo in Human Spaceflight Vehicles

An overview of more than 45 years of NASA human spaceflight experience is presented with respect to the thrust axis vibration response of liquid fueled rockets known as pogo. A coupled structure and propulsion system instability, pogo can result in the impairment of the astronaut crew, an unplanned engine shutdown, loss of mission, or structural failure. The NASA history begins with the Gemini Program and adaptation of the USAF Titan II ballistic missile as a spacecraft launch vehicle. It continues with the pogo experienced on several Apollo-Saturn flights in both the first and second stages of flight. The defining moment for NASA s subsequent treatment of pogo occurred with the near failure of the second stage on the ascent of the Apollo 13 mission. Since that time NASA has had a strict "no pogo" philosophy that was applied to the development of the Space Shuttle. The "no pogo" philosophy lead to the first vehicle designed to be pogo-free from the beginning and the first development of an engine with an integral pogo suppression system. Now, more than 30 years later, NASA is developing two new launch vehicles, the Ares I crew launch vehicle propelling the Orion crew excursion vehicle, and the Ares V cargo launch vehicle. A new generation of engineers must again exercise NASA s system engineering method for pogo mitigation during design, development and verification

A METHODOLOGY FOR CONDUCTING DESIGN TRADES FOR A SMALL SATELLITE LAUNCH VEHICLE WITH HYBRID ROCKET PROPULSION

The commercial space industry has recently seen a paradigm shift related to the launch of a small satellite into Low Earth Orbit. In the past, a small satellite was launched as a secondary payload with a medium or heavy launch vehicle where the primary payload placed a constraint on the orbit and schedule. Today, a dedicated launch of a small launch vehicle is the main operational concept to launch a small payload. Many Smallsat Launch Vehicles (SLV) have been under development by the commercial space industry to improve these launch services in recent years. Despite these efforts, the specific prices per launch are still high, and reducing these prices further remains a challenge. One promising technology candidate to reduce costs for SLV is hybrid rocket propulsion which has matured recently with some cost and safety advantages. Although hybrid rocket propulsion faces a number of challenges, including a low regression rate and combustion instabilities, academia and commercial companies have invested significant resources in developing this technology. With this motivation, this thesis has focused on the conceptual design of SLV with hybrid rocket propulsion. Moreover, a cost reduction strategy currently used by the commercial space industry was observed to be the development of a unique engine and using multiple of them in a launch vehicle. Following this trend, the vehicle concept investigated in this thesis was an expendable ground-launched vehicle with some architectural variables such as the number of stages and the number of hybrid motors in each stage. The design trade-off studies of such a small multistage launch vehicle with multiple hybrid motors in each stage require very long times especially when traditional point design approaches are used. As the number of design variables increase, the design space exploration becomes even more challenging. To provide a solution to this problem, a methodology for rapid conceptual design of such a vehicle was presented in this thesis. A physics-based conceptual design approach was followed in this study since SLV are relatively new concepts without much historical performance data. To conduct a multidisciplinary analysis, a physics-based, integrated modeling and simulation environment was constructed with four core disciplines: trajectory analysis, aerodynamics, propulsion, and weight. Aerodynamics and propulsion analysis were conducted using a first-principles approach, which was based on fundamental theories. A 3 Degree of Freedom (DOF) industrial, transparent, physics-based trajectory analysis software was used in this study based on availability. However, any other trajectory analysis software that a system designer is familiar with can be used in its place. In other words, the methodology developed in this thesis would remain unchanged if another trajectory analysis software were used. The weight discipline was represented at a high level by using Propellant Mass Fraction (PMF) design variable. A multidisciplinary modeling and simulation environment for launch vehicles may be computationally expensive depending on the fidelity levels of each discipline. Moreover, trajectory optimization is included in a launch vehicle design process conventionally which may be also computationally expensive depending on the optimization method. This expense poses a difficulty in performing a trade-off study for hundreds of vehicle design alternatives within the constraints of the schedule in the conceptual design phase. Because of this, trajectory optimization was removed from the design process to speed up the process by selecting a constant controller design. The methodology developed in this thesis consisted of two sequential steps. In the first step, a surrogate modeling approach was followed to replace the Modeling and Simulation (M&S) environment. A DOE method and a surrogate modeling method suitable to this problem were searched in this part. To cover the design space, a hybrid DOE consisting of a Fast Flexible Filling DOE and a three-level Full Factorial DOE was chosen. Artificial Neural Networks method was selected to fit approximation models because of the type of design variables (both continuous and discrete variables) and nonlinearity of the problem. The first experiment was conducted to test this hypothesis. As a result, it was demonstrated that this approach can provide accurate surrogate models for any desired response. In the second step, the specific mechanical energy-based design trade-off method was developed using some statistical methods. This method estimates the lower bound of the vehicles’ actual specific mechanical energy where the vehicles can be rapidly designed by using surrogate models. This lower bound was predicted with the help of the prediction interval of the specific mechanical energy’s model fit error. To fit the surrogate models, the necessary data were gathered by running the DOE in the integrated M&S environment while imposing some terminal conditions on the altitude of the vehicles analyzed in this environment. Specifically, the surrogate models of specific mechanical energy and flight path angle were used to design the vehicles rapidly. The second experiment was conducted to test this hypothesis. As a result, the actual specific mechanical energies computed via trajectory optimization were found to be consistent with the predictions. Overall, it was demonstrated that the proposed method enables a system designer to rapidly design some feasible vehicles, which can then proceed to the next design phase for further comparison, analysis, and design.M.S

Low cost propulsion systems for the developing world

Space has often been referred to as the final frontier. It is the curiosity of what lies beyond our planet that drives us to turn to the skies. This quest for knowledge and the chance of travelling to the heavens has compelled people to devote their lives to space science, innovation and analysis of our ever-expanding universe. Today the most significant impact of rocketry comes in the form of manned spaceflight. Vehicles like the Space Shuttle and Soyuz began the trend of greater commercialization of manned rocketry, enabling widespread access to space. Whilst the curiosity of what lies beyond may have propelled the development of the space tourism industry, its current operational cost is estimated as $20-$28 million per passenger per flight. Although the vision of providing low cost space travel still exists, its application is hindered by the costs associated with current space vehicles and mission operations. Furthermore, if we are to better understand our universe and are keen on commercializing space, we would require the space tourism industry to operate in a similar fashion to the aviation industry. As most current launch vehicles rely on chemical propulsion, the level of uncertainty in the market drives their fuel costs. In order to reduce the cost per flight, we must effectively increase the load factor per flight and operate multiple flights, enabling a greater number of paying passengers. In order to provide widespread access to space there needs to be a greater emphasis on the research and development of low cost Reusable Launch Vehicles (RLV) which predominantly rely on alternative fuel technologies, thereby reducing the overall cost per flight. Although progress would be slow, we would still be able to witness a boom in space tourism. This paper proposes the use of magnetic levitation and propulsion (Maglev) within a vacuum chamber as a viable low-cost propulsion technology. It aims to prove that such a system is capable of providing adequate thrust to future space vehicles. As Maglev systems allow for horizontal take-off and landing, such a launch system could be used in conjunction with current airports worldwide. Although the inception and creation of such a system may seem expensive, the long-term fiscal costs are relatively lower than current day systems. This is primarily because such a system relies on electrical power, whose supply and generation costs are much lower than that of chemical propellants. Also, the maintenance costs associated with the Maglev track are minimal, as during take-off there is no physical contact between the track and the launch vehicle. Similar to the aviation industry, the success of future space exploration programs and space tourism relies on international cooperation and alliances. This not only ensures that no one country dominates access to space, but also nurtures healthy competition by providing a level playing field. By implementing the afore mentioned system in politically stable developing nations, we ensure employment, innovation and motivation, all achieved through an international alliance. This system would not only ensure a faster urban development within these countries, but would also bring the vision of space science and exploration to a larger global audience. This paper discusses the overall cost analysis for a vacuum operated Maglev system, the various options available for the generation of power required by such a system and how the system’s long term costs can be aligned with the aviation industry

Velocity package Patent

High velocity guidance and spin stabilization gyro controlled jet reaction system for launch vehicle payload

Event-Driven Network Model for Space Mission Optimization with High-Thrust and Low-Thrust Spacecraft

Numerous high-thrust and low-thrust space propulsion technologies have been developed in the recent years with the goal of expanding space exploration capabilities; however, designing and optimizing a multi-mission campaign with both high-thrust and low-thrust propulsion options are challenging due to the coupling between logistics mission design and trajectory evaluation. Specifically, this computational burden arises because the deliverable mass fraction (i.e., final-to-initial mass ratio) and time of flight for low-thrust trajectories can can vary with the payload mass; thus, these trajectory metrics cannot be evaluated separately from the campaign-level mission design. To tackle this challenge, this paper develops a novel event-driven space logistics network optimization approach using mixed-integer linear programming for space campaign design. An example case of optimally designing a cislunar propellant supply chain to support multiple lunar surface access missions is used to demonstrate this new space logistics framework. The results are compared with an existing stochastic combinatorial formulation developed for incorporating low-thrust propulsion into space logistics design; our new approach provides superior results in terms of cost as well as utilization of the vehicle fleet. The event-driven space logistics network optimization method developed in this paper can trade off cost, time, and technology in an automated manner to optimally design space mission campaigns.Comment: 38 pages; 11 figures; Journal of Spacecraft and Rockets (Accepted); previous version presented at the AAS/AIAA Astrodynamics Specialist Conference, 201

Wind Tunnel Investigation of the Supersonic Stage Separation Aerodynamics of a Generic 0.0175-Scale Bimese Two-Stage-to-Orbit Reusable Launch Vehicle Configuration

A wind tunnel investigation was conducted of the supersonic stage separation aerodynamics of a generic two-stage-to-orbit bimese wingbody configuration in the NASA Langley Research Center Unitary Plan Wind Tunnel. Proximity and isolated model testing was conducted at Mach numbers of 2.3, 3.0, and 4.5 and a unit Reynolds number of 2.0 million per foot using 0.0175-scale models of the Langley Glide-Back Booster concept designated as the orbiter and booster in belly-to-belly and back-to-belly configurations. Longitudinal forces and moments were obtained on both models and surface static pressure measurements were obtained on the orbiter model at 328 relative proximity locations and at relative angles of attack of 0 degrees and 5 degrees. The test results supported a larger effort to develop and validate experimental and computational tools applicable to the design and simulation of stage separation and abort procedures for reusable launch vehicles composed of multiple bodies, including winged bodies. An initial proof-of-concept experiment featuring low-cost uninstrumented models was conducted to verify an emerging automated model control system and new support system hardware, and to identify potential model and support system blockage and unsteady aerodynamics/model dynamics prior to committing to higher-fidelity instrumented models. This investigation led to upgrades in the facility stage separation hardware, calibration and testing techniques and capabilities, and data analysis and documentation methodologies that have been extended to the more recent NASA Constellation and Space Launch System crew and cargo launch vehicle programs. A virtual diagnostics interface methodology was used to facilitate the design of the stage separation support hardware, to position the models in the test section, and to define the experimental test space. Advances in the facility automated model positioning system established a foundation for the development of a continuous-sweep data acquisition technique that is responsible for significant productivity improvements to the current NASA Space Launch System test program. The automated model positioning capability was leveraged to conduct a companion statistically-designed stage separation experiment requiring randomization of the relative proximity positions of the orbiter and booster models. The respective zones of influence and interference effects of the orbiter and booster were identified from three-dimensional scatter plots, contour and influence maps, and two-dimensional plotting methods. The highly-nonlinear, shock-dominated aerodynamic characteristics of the orbiter and booster in the Unitary Plan Wind Tunnel exhibited good agreement with independent test data obtained in a NASA Marshall Space Flight Center wind tunnel and with computational fluid dynamics predictions using a compressible, three-dimensional flow solver and an inviscid, unstructured Cartesian method

Prognostic Launch Vehicle Probability of Failure Assessment Methodology for Conceptual Systems Predicated on Human Causal Factors

Lessons learned from past failures of launch vehicle developments and operations were used to create a new method to predict the probability of failure of conceptual systems. Existing methods such as Probabilistic Risk Assessments and Human Risk Assessments were considered but found to be too cumbersome for this type of system-wide application for yet-to-be-flown vehicles. The basis for this methodology were historic databases of past failures, where it was determined that various faulty human-interactions were the predominant root causes of failure rather than deficient component reliabilities evaluated through statistical analysis. This methodology contains an expert scoring part which can be used in either a qualitative or a quantitative mode. The method produces two products: a numerical score of the probability of failure or guidance to program management on critical areas in need of increased focus to improve the probability of success. In order to evaluate the effectiveness of this new method, data from a concluded vehicle program (USAF's Titan IV with the Centaur G-Prime upper stage) was used as a test case. Although the theoretical vs. actual probability of failure was found to be in reasonable agreement (4.46% vs. 6.67% respectively) the underlying sub-root cause scoring had significant disparities attributable to significant organizational changes and acquisitions. Recommendations are made for future applications of this method to ongoing launch vehicle development programs

Exploiting technological synergies for future launch vehicles

Two launch vehicle concepts based on technologies available today or in a short term future in Western Europe are presented. The design of both launchers has the goal of exploiting synergies with current European programs to limit development and operational costs. Technologies of particular interest here are the high performance solid rocket motors with carbon-epoxy filament wound monolithic motor cases and the future high performance cryogenic expander cycle engine Vinci. The first concept dubbed ANGELA (A New GEneration LAuncher) is a study financed with funds of the German Ministry of Economics and managed by the DLR Space Administration. The project, which started in the summer of 2012 aims at designing a low cost versatile launcher able to place payloads between 2 and 5 tons into GTO. Three architectures have been considered during the first phase of the study. This phase was concluded in March 2013 with the preliminary stagings, which will be the starting point of more detailed analyses. The first architecture is made out of an H110 (stage with 110 tons of LOx/LH2) equipped with two Vulcain 2 engines with shortened nozzles and an H29 propelled by a Vinci engine. In addition the variation of the number of P36 solid rocket boosters allow to reach the entire range of payload performance. The second architecture differs from the first one only by the use of a new staged-combustion engine instead of two Vulcain 2 engines. The new engine, which should deliver 1800 kN in vacuum, allows a reduction of the size of the stages to H90-H24, enhanced with P34 boosters. The third and last architecture is a so called Multi PPH. The first stage is a bundle of 2 or 3 P120 solid rocket motors. The second stage is made out of one single P120, strictly similar to those used for the first stage. Finally the upper stage is an H23 equipped with a Vinci engine, the same as the two other architectures. The second launcher concept described in this paper is the small TSTO launch vehicle. It consists of a large solid rocket motor first stage P175 and a cryogenic upper stage propelled by the Vinci engine, H26. The preliminary design performed at DLR-SART considers two target performances. The light version of the small TSTO shall perform Galileo satellite replacement single launch missions to MTO corresponding to a payload performance of about 1400 kg in GTO. A heavy version of the launch vehicle shall be able to launch payloads up to 3000 kg in GTO. The performance increase for the heavy version is made possible by the addition of two pairs of P23 boosters, the second pair being ignited with a delay

Feasibility study of storage concepts for Scout and other NASA solid propellant launch vehicles

Feasibility study of long term storing of Scout and other solid propellant launch vehicles in assembled, flightworthy configuration and facility requirement