Advanced Hypervelocity Aerophysics Facility Workshop

The primary objective of the workshop was to obtain a critical assessment of a concept for a large, advanced hypervelocity ballistic range test facility powered by an electromagnetic launcher, which was proposed by the Langley Research Center. It was concluded that the subject large-scale facility was feasible and would provide the required ground-based capability for performing tests at entry flight conditions (velocity and density) on large, complex, instrumented models. It was also concluded that advances in remote measurement techniques and particularly onboard model instrumentation, light-weight model construction techniques, and model electromagnetic launcher (EML) systems must be made before any commitment for the construction of such a facility can be made

Numeric comparative study on Advanced Nozzles in subsonic counter-flows

openAl giorno d'oggi, la stragrande maggioranza dei razzi regolarmente commercializzati impiega convenzionali ugelli a campana, i quali si stanno avvicinando ai loro limiti di ottimizzazione. Alcuni concept di ugelli avanzati (ANC) offrono potenziali soluzioni alternative per la futura generazione di veicoli di lancio riutilizzabili (RLV). Nell'ottica migliorare le prestazioni delle manovre di recupero (Powered Descent and Landing), l'integrazione di queste tecnologie negli RLV sembra rivelarsi promettente, grazie alla loro capacità di auto-adattamento al variare dell'altitudine. In particolare, durante la fase finale della manovra -- nota come landing burn -- il veicolo affronta una fase di retro-propulsione subsonica, in cui il motore viene acceso in direzione opposta al contro-flusso ambientale. Questo studio mira a valutare e confrontare le prestazioni aerodinamiche e degli ugelli di un RLV durante quest'ultima fase. Ciò è condotto nell'ecosistema del software ANSYS Fluent, tramite simulazioni CFD di quattro distinti concept di ugello: convenzionale ugello a campana (due diversi design: un profilo Rao parabolico e un Truncated-Ideal-Contour), Aerospike, Expansion-Deflection e Dual-Bell. Ognuna di queste tipologie è sottoposta a sua volta a test in quattro scenari: static burn on- e off-design (rispettivamente motore funzionante al punto di progetto e a SLS, senza alcun controflusso), discesa aerodinamica (con motore spento ma contro-flusso attivato) e retro-propulsione subsonica (con motore e controflusso attivati). I risultati numerici vengono successivamente confrontati e convalidati con una parallela campagna sperimentale condotta sugli stessi casi in laboratorio.In the present era, the vast majority of regularly manufactured rockets employ conventional Bell Nozzles, which are approaching their limits of improvement. Advanced Nozzle Concepts (ANCs) offer potential alternative solutions for the future generation of Reusable Launch Vehicles (RLVs). Integrating these technologies into RLVs holds promise for enhancing the performance of Powered Descent and Landing (PDL) recovery maneuvers, owing to their capacity for altitude compensation. Specifically, during the final stage of PDL -- known as the landing burn maneuver -- the vehicle undergoes a subsonic retro-propulsion phase, wherein the engine is ignited against a low-speed counter-flow. This study aims to assess and compare the aerodynamic and nozzle performance of a reusable launch vehicle during this latter phase. The evaluation is conducted using the ANSYS Fluent software environment, employing Computational Fluid Dynamics (CFD) numerical simulations of four distinct nozzle concepts: conventional bell nozzle (including two different designs: Rao parabolic and Truncated-Ideal-Contour), Aerospike, Expansion-Deflection, and Dual Bell nozzles. Each of these concepts is subjected to testing under four scenarios: on- and off-design static burns (respectively with the engine operating at the design point and at SLS, without any counter-flow), aerodynamic descent (with the engine off and counter-flow activated), and subsonic retro-propulsion (with both the engine and counter-flow activated). The numerical results are subsequently compared and validated against a parallel experimental campaign conducted on the same cases

Scaled Rocket Testing in Hypersonic Flow

NASA's Space Launch System (SLS) uses four clustered liquid rocket engines along with two solid rocket boosters. The interaction between all six rocket exhaust plumes will produce a complex and severe thermal environment in the base of the vehicle. This work focuses on a recent 2% scale, hot-fire SLS base heating test. These base heating tests are short-duration tests executed with chamber pressures near the full-scale values with gaseous hydrogen/oxygen engines and RSRMV analogous solid propellant motors. The LENS II shock tunnel/Ludwieg tube tunnel was used at or near flight duplicated conditions up to Mach 5. Model development was strongly based on the Space Shuttle base heating tests with several improvements including doubling of the maximum chamber pressures and duplication of freestream conditions. Detailed base heating results are outside of the scope of the current work, rather test methodology and techniques are presented along with broader applicability toward scaled rocket testing in supersonic and hypersonic flow

Human Mars Entry, Descent and Landing Architecture Study: Deployable Decelerators

NASAs Entry, Descent and Landing Architecture Study uses a trajectory simulation framework to evaluate various technologies and concepts of operations for human scale EDL at Mars. The study results inform agency technology investments. This paper summarizes the design assumptions and analysis of two deployable entry concepts performed in Phase 2 of the study. The entry concepts include a rigid deployable called the Adaptable Deployable Entry Placement Technology and an inflatable concept called the Hypersonic Inflatable Aerodynamic Decelerator. This paper describes the concept operations of these vehicles to deliver a 20-metric ton payload to the surface of Mars. Details of vehicle design and flight performance are summarized along with results of analysis on the aft body heating and its effect on the payload. Finally, recommended technology investments based on the results are presented

Propulsion Controls and Diagnostics Research in Support of NASA Aeronautics and Exploration Mission Programs

The Controls and Dynamics Branch (CDB) at National Aeronautics and Space Administration (NASA) Glenn Research Center (GRC) in Cleveland, Ohio, is leading and participating in various projects in partnership with other organizations within GRC and across NASA, the U.S. aerospace industry, and academia to develop advanced propulsion controls and diagnostics technologies that will help meet the challenging goals of NASA programs under the Aeronautics Research and Exploration Systems Missions. This paper provides a brief overview of the various CDB tasks in support of the NASA programs. The programmatic structure of the CDB activities is described along with a brief overview of each of the CDB tasks including research objectives, technical challenges, and recent accomplishments. These tasks include active control of propulsion system components, intelligent propulsion diagnostics and control for reliable fault identification and accommodation, distributed engine control, and investigations into unsteady propulsion systems

Experiments and simulations of hybrid rocket internal flows and material behaviour

In the last decade a significant and ever growing interest has been addressed towards hybrid rocket propulsion, which offers the best-of-both-worlds by leveraging the favourable aspect of both traditional solid and liquid systems. Among the numerous advantages which characterize hybrid rockets, the most attractive ones are the re-ignition and throttling capabilities combined with the possibility of embedding environmentally sustainable propellants and, of the utmost importance, their intrinsic safety and lower operational costs. Moreover, hybrid rockets yield a better specific impulse than solid propellant rockets and a higher density impulse than liquids, which make them a promising technology in a number of space missions. The widely recognized potentialities of the hybrid rocket warrant the renewed research efforts that are being devoted to its development, but the state-of-the-art of this technology still presents a number of challenging issues to be solved. A first fundamental task is the definition of suitable models for the prediction of the motor internal ballistics and performance. In particular, rocket performance is governed by the rate at which the fuel is gasified, i.e. by the fuel regression rate, as this latter determines the total mass flow rate and the overall oxidizer-to-fuel mixture ratio, which, for a given chamber pressure, control the motor thrust and the ideal specific impulse. For a given fuel, regression rate is basically limited by the heat flux input to the solid grain, which mainly depends on the thermo-fluid-dynamics in the combustion chamber. This latter is significantly influenced by several geometrical parameters, such as, for example, the oxidizer injection configuration or the grain port shape. Furthermore, the recent efforts aimed at overcoming the main drawback of the hybrid rockets, which is the low regression rate of conventional polymeric fuels, have been focused on the development of new paraffin-based fuels, characterized by a consumption mechanism presenting additional complex phenomena compared to that of conventional polymers. Their intrinsic characteristic is the onset of a thin liquid layer on the fuel grain surface, which may become unstable, leading to the lift-off and entrainment of fuel liquid droplets into the main gas stream, increasing the fuel mass transfer rate. This phenomenon is strongly susceptible to the fuel composition, its manufacturing process and the obtained thermo-mechanical properties as well as to the engine operating conditions, which makes the prediction of the regression rate and combustion chamber internal ballistics even harder than in the case of a pure polymer. In this framework, computational fluid dynamics of hybrid rocket internal ballistics is becoming a key tool for reducing the engine operation uncertainties and development cost, but its application still presents numerous challenges due to the complexity of modelling the phenomena involved in the fuel consumption mechanism and the interaction with the reacting flowfield, for both the cases of classical polymeric and liquefying paraffin-based fuels. A research effort is therefore of major importance in order to cover the lacking aspects and obtain quantitatively accurate results. Another challenge for the hybrid rocket technology development is the optimization of the design of thermal insulations. The inner surface of the exhaust nozzle, through which the flow is accelerated to supersonic conditions producing the required thrust, is the most critical in this sense, as it is subjected to the highest shear stress and heat fluxes in a chemically aggressive environment. These severe conditions usually lead to removal of surface material due to heterogeneous reactions between oxidizing species in the hot gas and the solid wall. Because of the material erosion, there is an enlargement of the nozzle throat section and a consequent decrease of rocket thrust, with detrimental effects over the motor operation. Thus, the requirement that dimensional stability of the nozzle throat should be maintained makes the selection of suitable rocket nozzle materials extremely hard. In recent years, Ultra-High-Temperature Ceramics (UHTC) and Ultra-High-Temperature Ceramic Matrix Composites are the subject of considerable interest as innovative materials for rocket application, but still need to be properly characterized. Experimental testing along with computational fluid dynamic (CFD) simulations are, thus, both needed to improve the design and the current performance prediction capabilities of such propulsion systems. In this framework, the University of Naples is involved in the European project C3HARME – Next Generation Ceramic Composites for Combustion Harsh Environment and Space, in collaboration with other research centres, universities and industries, which aims at the design, manufacturing and testing of new-class high-performance UHTCMC for near-zero erosion rocket nozzles. In the present work, the above-mentioned challenges are dealt with taking a combined experimental/numerical approach to improve understanding of the interaction between the gaseous combusting flow typical of hybrid rocket engines and the surface of solid materials involved in their operation, with a special focus to the fuel grain present in the combustion chamber, with the aim of predicting its consumption mechanism, and the exhaust nozzle inner surface, with the aim of identifying and validating new-class UHTCMC materials with improved erosion and structural resistance to the severe conditions experienced in particular in the throat region. In particular, the first main objective of the present work is the definition of proper computational thermo-fluid-dynamic models of the hybrid rocket internal ballistics, including a dedicated gas/surface interface treatment based on local mass, energy and mean mixture fraction balances as well as proper turbulence boundary conditions, which can properly model the physical fuel consumption mechanism in both the cases of polymeric and liquefying fuels. For the validation of the computational models, a number of experimental test cases, obtained from static firing of laboratory scale rockets, have been performed at the Aerospace Propulsion Laboratory of University of Naples “Federico II” and successively numerically reconstructed. The comparison between the numerical results and the corresponding experimental data allowed validating the adopted model and identifying possible future improvements. Then, the research activities for the characterization of new-class UHTCMC materials are presented and discussed. This part of the work was mainly focused on an extensive experimental campaign for the characterization of new-class UHTCMC materials. In particular, first preliminary tests on small samples exposed to the supersonic exhaust jet of a 200N-class hybrid rocket operated with gaseous oxygen burning cylindrical port High-Density PolyEthylene (HDPE) fuel grains have been carried out for a fast characterization and a preliminary screening of the best candidates for the final applications. After that UHTCMC nozzle throat inserts has been manufactured and experimentally tested to verify the erosion resistance and evaluate the effects on the rocket performance by comparison with those obtained in similar operating conditions employing a graphite nozzle. The experimental activities are supported by simplified low-computational-cost numerical simulations, whose main objectives has been the prediction of the complex flow field in the hybrid rocket combustion chamber and the thermo-fluid dynamic conditions on the material. Future research activities will be then focused to the further development of the numerical models with the extension of the treatment for the gaseous flow/solid surface interaction in order to get a deeper insight on the new materials behaviour

A Survey of Gaps, Obstacles, and Technical Challenges for Hypersonic Applications

The object of this study is to canvas the literature for the purpose of identifying and compiling a list of Gaps, Obstacles, and Technological Challenges in Hypersonic Applications (GOTCHA). The significance of GOTCHA related deficiencies is discussed along with potential solutions, promising approaches, and feasible remedies that may be considered by engineers in pursuit of next generation hypersonic vehicle designs and optimizations. Based on the synthesis of several modern surveys and public reports, a cohesive list is formed, consisting of widely accepted areas needing improvement and falling under several general categories. These include: aerodynamics, propulsion, materials, analytical modeling, CFD modeling, and education in high speed flow physics. New methods and lines of research inquiries are suggested such as the homotopy-based analysis (HAM) for the treatment of strong nonlinearities, the use of improved turbulence models and unstructured grids in numerical simulations, the need for accessible validation data, and the refinement of mission objectives for Hypersonic Air-Breathing Propulsion (HABP)

Research reports: 1991 NASA/ASEE Summer Faculty Fellowship Program

The basic objectives of the programs, which are in the 28th year of operation nationally, are: (1) to further the professional knowledge of qualified engineering and science faculty members; (2) to stimulate an exchange of ideas between participants and NASA; (3) to enrich and refresh the research and teaching activities of the participants' institutions; and (4) to contribute to the research objectives of the NASA Centers. The faculty fellows spent 10 weeks at MSFC engaged in a research project compatible with their interests and background and worked in collaboration with a NASA/MSFC colleague. This is a compilation of their research reports for summer 1991

Compilation of Abstracts for SC12 Conference Proceedings

1 A Breakthrough in Rotorcraft Prediction Accuracy Using Detached Eddy Simulation; 2 Adjoint-Based Design for Complex Aerospace Configurations; 3 Simulating Hypersonic Turbulent Combustion for Future Aircraft; 4 From a Roar to a Whisper: Making Modern Aircraft Quieter; 5 Modeling of Extended Formation Flight on High-Performance Computers; 6 Supersonic Retropropulsion for Mars Entry; 7 Validating Water Spray Simulation Models for the SLS Launch Environment; 8 Simulating Moving Valves for Space Launch System Liquid Engines; 9 Innovative Simulations for Modeling the SLS Solid Rocket Booster Ignition; 10 Solid Rocket Booster Ignition Overpressure Simulations for the Space Launch System; 11 CFD Simulations to Support the Next Generation of Launch Pads; 12 Modeling and Simulation Support for NASA's Next-Generation Space Launch System; 13 Simulating Planetary Entry Environments for Space Exploration Vehicles; 14 NASA Center for Climate Simulation Highlights; 15 Ultrascale Climate Data Visualization and Analysis; 16 NASA Climate Simulations and Observations for the IPCC and Beyond; 17 Next-Generation Climate Data Services: MERRA Analytics; 18 Recent Advances in High-Resolution Global Atmospheric Modeling; 19 Causes and Consequences of Turbulence in the Earths Protective Shield; 20 NASA Earth Exchange (NEX): A Collaborative Supercomputing Platform; 21 Powering Deep Space Missions: Thermoelectric Properties of Complex Materials; 22 Meeting NASA's High-End Computing Goals Through Innovation; 23 Continuous Enhancements to the Pleiades Supercomputer for Maximum Uptime; 24 Live Demonstrations of 100-Gbps File Transfers Across LANs and WANs; 25 Untangling the Computing Landscape for Climate Simulations; 26 Simulating Galaxies and the Universe; 27 The Mysterious Origin of Stellar Masses; 28 Hot-Plasma Geysers on the Sun; 29 Turbulent Life of Kepler Stars; 30 Modeling Weather on the Sun; 31 Weather on Mars: The Meteorology of Gale Crater; 32 Enhancing Performance of NASAs High-End Computing Applications; 33 Designing Curiosity's Perfect Landing on Mars; 34 The Search Continues: Kepler's Quest for Habitable Earth-Sized Planets

Pluto Hop, Skip, and Jump

This is the Final Report from Global Aerospace Corporation on this NIAC effort (Grant Nos.: NNX17AJ71G and 80NSSC18K0062) to develop the Pluto Hop, Skip, and Jump mission concept. We sought out to establish the feasibility of using a large inflatable drag device to decelerate and land on Pluto from interplanetary speed (~14 km/s) using only the Pluto atmosphere and just a few kilograms of propellant. The design and analysis efforts in Phase I indicated that this is feasible. Aerodynamic heating and loads were found to be orders of magnitude less than typical planetary entries due to the ultra-low ballistic coefficient craft and the low density and large scale height of the Pluto atmosphere. The deceleration system is capable of delivering a 200-kg lander-hopper to the surface or inserting an orbiter of a similar mass using aerocapture. Mission analysis work led to a reference mission with Earth launch in 2029, Jupiter assist in 2030, and Pluto arrival in 2040