In this paper, we will examine a configuration for a reusable military launch vehicle (RMLS) concept. This configuration allows for the vehicle to land in an inverted attitude. Such inverted landing improves the turnaround time of the vehicle by reducing the maintenance requirements of the vehicle's thermal protection system. An analysis is performed to examine the impacts by the configuration on stability, control, and footprint for an RMLS configuration.

The ability to compute the maximum area on the earth's surface (footprint) reachable by an autonomous air vehicle can be useful in planning for the vehicle's safe operations. The information can be important when the vehicle experiences subsystem failures causing it to be unable to maintain its nominal performance. In this paper, we present a method to calculate the footprint of a reusable launch vehicle that experiences a failure in one or more of its aero-control surfaces. During a control effector failure, the maximum attainable moments of the vehicle are reduced, which may decrease the range of conditions that the vehicle can maintain a trimmed condition. Additionally, the lift and drag characteristics of the vehicle can change when control effectors are moved to off-nominal positions to correct for moment imbalance caused by failures or damage. As a result, the footprint of the vehicle is reduced. A technique for calculating the available effectiveness of the aero-control surfaces is used in conjunction with a footprint generation algorithm to include the effects of rotational trim on the vehicle footprint.

The key to opening the use of space to private enterprise and to broader public uses lies in reducing the cost of the transportation to space. More routine, affordable access to space will entail aircraft-like quick turnaround and reliable operations. Currently, the space Shuttle is the only reusable launch vehicle, and even parts of it are expendable while other parts require frequent and extensive refurbishment. NASA's highest priority new activity, the Reusable Launch Vehicle program, is directed toward developing technologies to enable a new generation of space launchers, perhaps but not necessarily with single stage to orbit capability. This book assesses whether the technology development, test and analysis programs in propulsion and materials-related technologies are properly constituted to provide the information required to support a December 1996 decision to build the X-33, a technology demonstrator vehicle; and suggest, as appropriate, necessary changes in these programs to ensure that they will support vehicle feasibility goals.

This thesis considers the problem of generating optimal entry trajectories for a reusable launch vehicle following a control surface failure. The thesis builds upon the work of Dr. David Doman, Dr. Michael Oppenheimer and Dr. Michael Bolender of the Air Vehicles Directorate, Air Force Research Lab Dayton Ohio. The primary focus of this work is to demonstrate the feasibility of inner loop reconfiguration and outer loop trajectory retargeting and replanning for the X-33 reusable launch vehicle (RLV) following the imposition of a control surface failure. The trajectory generation model employs path constraints generated by an AFRL trim deficiency algorithm coupled with an inner loop control allocator and aerodynamic database that captures the full 6-DOF vehicle aerodynamic effects while utilizing an outer loop 3-DOF model. The resulting optimal trajectory does not violate the trim deficiency constraints and provides additional margins for trajectories flown during failure conditions. The footprints generated by the thesis show that contemporary footprint analysis for vehicles experiencing control surface failures are overly optimistic when compared to those footprints that consider vehicle aerodynamic stability and realistic landable attitudes at the threshold of the landing runway. The results of the thesis also show the performance reductions resulting from decoupling the inner and outer loop and that trajectories can be generated to the landing runway without using a region of terminal area energy management.

The Space Launch Initiative (SLI) program is developing a second-generation reusable launch vehicle. The program goals include lowering the risk of loss of crew to 1 in 10,000 and reducing annual operations cost to one third of the cost of the Space Shuttle. The SLI missions include NASA, military and commercial satellite launches and crew and cargo launches to the space station. The SLI operations analyses provide an assessment of the operational support and infrastructure needed to operate candidate system architectures. Measures of the operability are estimated (i.e. system dependability, responsiveness, and efficiency). Operations analysis is used to determine the impact of specific technologies on operations. A conceptual path to reducing annual operations costs by two thirds is based on key design characteristics, such as reusability, and improved processes lowering labor costs. New operations risks can be expected to emerge. They can be mitigated with effective risk management with careful identification, assignment, tracking, and closure. SLI design characteristics such as nearly full reusability, high reliability, advanced automation, and lowered maintenance and servicing coupled with improved processes are contributors to operability and large operating cost reductions.

Realizing a reusable launch vehicle (RLV) that is low cost with highly effective launch capability has become the "Holy Grail" within the aerospace community world-wide. Clear understanding of the vehicle's operational limitations and flight characteristics in all phases of the flight are preponderant components in developing such a launch system. This dissertation focuses on characterizing and designing the RLV optimal trajectories in order to aid in strategic decision making during mission planning in four areas: 1) nominal ascent phase, 2) abort scenarios and trajectories during ascent phase including abort-to-orbit (ATO), transoceanic-abort-landing (TAL) and return-to-launch-site (RTLS), 3) entry phase (including footprint), and 4) systems engineering aspects of such flight trajectory design. The vehicle chosen for this study is the Lockheed Martin X-33 lifting-body design that lifts off vertically with two linear aerospike rocket engines and lands horizontally. An in-depth investigation of the optimal endo-atmospheric ascent guidance parameters such as earliest abort time, engine throttle setting, number of flight phases, flight characteristics and structural design limitations will be performed and analyzed to establish a set of benchmarks for making better trade-off decisions. Parametric analysis of the entry guidance will also be investigated to allow the trajectory designer to pinpoint relevant parameters and to generate optimal constrained trajectories. Optimal ascent and entry trajectories will be generated using a direct transcription method to cast the optimal control problem as a nonlinear programming problem. The solution to the sparse nonlinear programming problem is then solved using sequential quadratic programming. Finally, guidance system hierarchy studies such as work breakdown structure, functional analysis, fault-tree analysis, and configuration management will be developed to ensure that the guidance system meets the definition of vehicle design requirements and constraints.