Jeffrey V. Bowles

Near-optimal propulsion-system operation for an air-breathing launch vehicle

A methodology for determining the near-optimal operation of the propulsion system of hybrid air-breathing launch vehicles is derived. The method is based on selecting propulsion-system modes and parameters that maximize a certain performance function. This function is derived from consideration of the energy-state model of the aircraft equations of motion. The vehicle model reflects the many interactions and complexities of the multimode air-breathing and rocket engine systems proposed for launch-vehicle use. The method is used to investigate the optimal throttle switching of air-breathing and rocket engine modes, and to investigate the desirability of using liquid-oxygen augmentation in air-breathing engine cycles, the oxygen either carried from takeoff or collected in flight. It is found that the air-breathing engine is always at full throttle, and that the rocket is on full at takeoff and at very high Mach numbers, but off otherwise. Augmentation of the air-breathing engine with stored liquid oxygen is beneficial, but only marginally so.

Optimal trajectories for hypersonic launch vehicles

In this paper, we derive a near-optimal guidance law for the ascent trajectory from earth surface to earth orbit of a hypersonic, dual-mode propulsion, lifting vehicle. Of interest are both the optimal flight path and the optimal operation of the propulsion system. The guidance law is developed from the energy-state approximation of the equations of motion. Because liquid hydrogen fueled hypersonic aircraft are volume sensitive, as well as weight sensitive, the cost functional is a weighted sum of fuel mass and volume; the weighting factor is chosen to minimize gross take-off weight for a given payload mass and volume in orbit.

TPS SELECTION AND SIZING TOOL IMPLEMENTED IN AN ADVANCED ENGINEERING ENVIRONMENT

A tool, TPSSIZER, was developed to provide a Thermal Protection System (TPS) analysis and design capability . The tool focuse d on analysis of space vehicles at the conceptual design level and was implemented in a collaborative engineering analysis environment. TPS sizing methodologies and data exchange interfaces with supporting disciplines were developed. Additionally, improv ements were made to prior art by introducing automatic generation of TPS stackups, automatic generation of aerothermal environment files, maintenance of consistent material properties descriptions, and the capability to simultaneously ev aluate multiple nom inal and abort flight trajectories .

Minimum Heating Entry Trajectories for Reusable Launch Vehicles

A e nite control volume heat transfer analysis is coupled to a e ight-path optimization and integration algorithm forthepurposeofcalculatingconductiveheatratesand transienttemperatureeffectswithin thethermalprotection system of a reusable launch vehicle. Results are obtained for three different thermal protection system concepts: tile, blanket, and metallic. The optimization algorithm is based on the energy state approximation and is used to generate optimal entry trajectories minimizing the following three criteria: 1 ) the thermal energy absorbed at the vehicle surface, 2 ) the heat load applied to the vehicle, and 3 ) the thermal energy absorbed by the internal structure. Results indicate that allthreetrajectoriesproduce comparablepeak internal structure temperatures for a given thermal protection system, with the trajectory minimizing the heat load applied to the vehicle producing thelowestpeaktemperature.However,ifthemaximum stagnation temperatureconstraintatthenoseof thevehicle is increased from 3000 to 4000 ±F, the trajectory minimizing the thermal energy absorbed by the internal structure becomes superior. Further, the trajectory with the 4000 ±F limit gives a peak internal structure temperature 25 ±F less than the one with the 3000 ± F limit.

Trajectory and Thermal Protection System Design for Reusable Launch Vehicles

Geometry, aero/aerothermal, trajectory, and thermal protection selection and sizing tools are linked together in a collaborative engineering environment to form a multidisciplinary analysis model of a reusable launch vehicle performing atmospheric entry. This entry model can determine the effects of vehicle shape and trajectory on the aerothermal environment that must be endured by the thermal protection system. The aerothermal environment in turn determines the size and weight of the thermal protection system required by the vehicle. The importance of interdisciplinary coupling on the design of the vehicle thermal protection system is demonstrated for a wing-body vehicle. Results of a parametric study of the influence of wing thickness on maximum cross-range and thermal protection weight are reported. Difficulties encountered with trajectory optimization are discussed.

Mars Sample Return using commercial capabilities: Mission architecture overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for a MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. The major element required for the MSR mission are described and include an integration of the emerging commercial capabilities with small spacecraft design techniques; new utilizations of traditional aerospace technologies; and recent technological developments. We report the feasibility of a complete and closed MSR mission design using the following scenario that can start in any one of three Earth to Mars launch opportunities, beginning in 2022: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV); an Earth Return Vehicle (ERV); and hardware to transfer a sample collected in a previously landed rover mission, such as the Mars 2020 rover, to the ERV. The Red Dragon descends to land on the surface of Mars using Supersonic Retro Propulsion (SRP). After previously collected samples are transferred to the ERV, the single-stage MA V launches the ERV from the surface of Mars to a Mars phasing orbit. The MA V uses a storable liquid, pump-fed bi-propellant propulsion system. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEl) burn. Once near Earth the ERV performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit (LTO)an Earth orbit, at lunar distance. A later mission, using a Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft, makes a controlled Earth re-entry preventing any unintended release of pristine Martian materials into the Earths biosphere. Other capsule type vehicles and associated launchers may be applicable. An MSR launch in 2022 becomes the preferred option if the Mars 2020 rover is the previous sample caching vehicle. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MA V exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX; published analyses from other sources; as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance were derived. An assessment of the entry condition effects on the thermal protection system (TPS), currently in use for Dragon missions, showed no significant stressors. A useful mass of 2.0 mt is provided and includes mass growth allowances for the MA V, the ERV, and mission unique equipment. We also report on alternate propellant options for the MA V and options for the ERV, including propulsion systems; crewed versus robotic retrieval mission; as well as direct Earth entry. International Planetary Protection (PP) policies as well as verifiable means of compliance with both forward and back contamination controls, will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. This work shows that emerging commercial capabilities can be effectively integrated into a mission to achieve an important planetary science objective.

An Overview of the Role of Systems Analysis in NASA's Hypersonics Project

NASAs Aeronautics Research Mission Directorate recently restructured its Vehicle Systems Program, refocusing it towards understanding the fundamental physics that govern flight in all speed regimes. Now called the Fundamental Aeronautics Program, it is comprised of four new projects, Subsonic Fixed Wing, Subsonic Rotary Wing, Supersonics, and Hypersonics. The Aeronautics Research Mission Directorate has charged the Hypersonics Project with having a basic understanding of all systems that travel at hypersonic speeds within the Earths and other planets atmospheres. This includes both powered and unpowered systems, such as re-entry vehicles and vehicles powered by rocket or airbreathing propulsion that cruise in and accelerate through the atmosphere. The primary objective of the Hypersonics Project is to develop physics-based predictive tools that enable the design, analysis and optimization of such systems. The Hypersonics Project charges the systems analysis discipline team with providing it the decision-making information it needs to properly guide research and technology development. Credible, rapid, and robust multi-disciplinary system analysis processes and design tools are required in order to generate this information. To this end, the principal challenges for the systems analysis team are the introduction of high fidelity physics into the analysis process and integration into a design environment, quantification of design uncertainty through the use of probabilistic methods, reduction in design cycle time, and the development and implementation of robust processes and tools enabling a wide design space and associated technology assessment capability. This paper will discuss the roles and responsibilities of the systems analysis discipline team within the Hypersonics Project as well as the tools, methods, processes, and approach that the team will undertake in order to perform its project designated functions.

A Rigid Mid Lift-to-Drag Ratio Approach to Human Mars Entry, Descent, and Landing

Current NASA Human Mars architectures require delivery of approximately 20 metric tons of cargo to the surface in a single landing. A proposed vehicle type for performing the entry, descent, and landing at Mars associated with this architecture is a rigid, enclosed, elongated lifting body shape that provides a higher lift-to-drag ratio (L/D) than a typical entry capsule, but lower than a typical winged entry vehicle (such as the Space Shuttle Orbiter). A rigid Mid-L/D shape has advantages for large mass Mars EDL, including loads management, range capability during entry, and human spaceflight heritage. Previous large mass Mars studies have focused more on symmetric and/or circular cross-section Mid-L/D shapes such as the ellipsled. More recent work has shown performance advantages for non-circular cross section shapes. This paper will describe efforts to design a rigid Mid-L/D entry vehicle for Mars which shows mass and performance improvements over previous Mid-L/D studies. The proposed concept, work to date and evolution, forward path, and suggested future strategy are described.

Skylon Aerodynamics and SABRE Plumes

*† ‡ § An independent partial assessment is provided of the technical viability of the Skylon aerospace plane concept, developed by Reaction Engines Limited (REL). The objectives are to verify REL’s engineering estimates of airframe aerodynamics during powered flight and to assess the impact of Synergetic Air-Breathing Rocket Engine (SABRE) plumes on the aft fuselage. Pressure lift and drag coefficients derived from simulations conducted with Euler equations for unpowered flight compare very well with those REL computed with engineering methods. The REL coefficients for powered flight are increasingly less acceptable as the freestream Mach number is increased beyond 8.5, because the engineering estimates did not account for the increasing favorable (in terms of drag and lift coefficients) effect of underexpanded rocket engine plumes on the aft fuselage. At Mach numbers greater than 8.5, the thermal environment around the aft fuselage is a known unknown−a potential design and/or performance risk issue. The adverse effects of shock waves on the aft fuselage and plumeinduced flow separation are other potential risks. The development of an operational reusable launcher from the Skylon concept necessitates the judicious use of a combination of engineering methods, advanced methods based on required physics or analytical fidelity, test data, and independent assessments.

Conceptual stage separation from widebody subsonic carrier aircraft for space access

Stage separation is a critical technical issue for developing two-stage-to-orbit (TSTO) launch systems with widebody carrier aircraft that use air-breathing propulsion and launch vehicle stages that use rocket propulsion. During conceptual design phases, this issue can be addressed with a combination of engineering methods, computational fluid dynamics simulations, and trajectory analysis of the mated system and the launch vehicle after staging. The outcome of such analyses helps to establish the credibility of the proposed TSTO system and formulate a ground-based test programme for the preliminary design phase. This approach is demonstrated with an assessment of stage separation from the shuttle carrier aircraft. Flight conditions are determined for safe mated flight, safe stage separation, and for the launch vehicle as it commences ascending flight. Accurate assessment of aerodynamic forces and moments is critical during staging to account for interference effects from the proximities of the two large vehicles. Interference aerodynamics have a modest impact on the separation conditions and separated flight trajectories, but have a significant impact on the interaction forces.