INTRODUCTION to the WORKSHOP Dr. G. Ortega 2021, May 11th
ESA UNCLASSIFIED - For Official Use
1 …the process of technology
1 I have a great idea!
2 I write a proposal
3 My proposal is approved
4 I execute the activity of the proposal
2 Welcome !
•Original idea of the TEC-MPA Section of ESA and materialized in 2021 •Is there any interest on Will the idea of the discussion ideas for technology pre-discussions work? prior we make any kind of proposal? •Number of registrations: 161
3 Objectives
• Objective 1: Provide for an overview of the current state of technology ideas • Objective 2: Gather inputs from possible Bidders • Objective 3: Prioritize the upcoming research and development inputs to the ESA technology plans (TRP, GSTP) for the coming cycle
4 Scope of the Workshop
• 1) Technology for the architecture design, analysis and technical assessment of space transportation vehicles for suborbital, orbital and exploration applications, including upper stages, (re)-entry, expendable, and reusable vehicles • 2) Technology for the development of tools and techniques for the feasibility and viability assessments, and quick design iterations of flight vehicles • 3) Technology for flight vehicles, flight physics, aerodynamics, thermodynamics and fluid dynamics engineering and the architecture design and analysis of suborbital, (re-)entry, space transportation, and exploration vehicles
5 Time Line
1 2 3 4 5
May June September October November
Workshop Consolidation of Final list of the best Preparation of Introduction of the ideas ideas (including upcoming TDE ideas in ESA description) cycle internal system
6 The program
7 Can I take your idea and ….?
•No, please •The ideas are provided to you with the aim of discussion •The information should not be used to move the ideas to non-ESA technology programs
8 I have another idea…
•Can I please discuss with you privately? •Yes, of course. No problem. See points of contact
9 ESA points of contact
Csaba Jeroen Johan Luca
Louis Orr Richard Stephan Victor
10 Thank you to the organizers!
Csaba
Orr Stephan
11 Have a great Workshop!
ESA UNCLASSIFIED - For Official Use
12 Day 1
13 Prediction Fluid Dynamics Methods for Propellant Tanks
Prog. GSTP Budget: 800K Duration: 24 months Aim: A Ref.: A 04 TRL: 4 to 5
Objective(s): To develop prediction methods for cryogenic multi-phase flows to application level and provide experimental data from breadboard, on ground micro gravity facilities and flight experiments that can be utilized to validate such methods. Quantification of margins of uncertainty for experimental and numerical data shall be investigated, in detail.
Description Numerical techniques are developed and validated in a structured manner for the different relevant environments from ground conditions to micro gravity and isothermal single phase applications to non-isothermal with phase changes. Incremental validation is applied, i.e. all new levels of modelling have to demonstrate that all lower levels of required modelling and complexity can still be treated successfully. Resulting CFD tools are required to be computationally efficient, empirical models shall be proposed to avoid excessive computational effort for flow accurate prediction of phenomena that require very high temporal spatial resolution of the applied CFD method. Quantifiable margins of uncertainty as consequence of incomplete modelling have to be accepted to allow calculation of flow configurations that are typical for todays applications.
Background and previous activities Long history of CFD development and related validation testing.
Technical Officer: Richard Schwane ([email protected]) 14 Interaction of Fluid and Structures under Sloshing Conditions
Prog. GSTP Budget 800K Duration: 24 months Aim: A Ref.: A 05 TRL: 4 to 5
Objective(s) This activity shall assess functionality of FSI methods for the prediction of fluid motion and structure response, resulting forces on structures and effect on space craft trajectory and attitude. Storable propellants, commonly used in satellites and cryogenic propellants used on launch vehicles shall be considered. Material characteristics of tanks and support frames shall be analyzed for cryogenic applications. To prepare refueling of cryogenic tanks in orbit, also flow kinematics of valves, bubble pressure grids and vanes shall be investigated.
Description Sloshing creates forces on structure elements that are transmitted to the spacecraft. Such forces have an effect on spacecraft dynamics and have to be known by the GNC system to provide commands to attitude control systems to keep the spacecraft in a stable and controlled condition. Material properties, including materials that might show advantages for the specific requirements of fuel reservoirs in orbit shall be investigated in detail, considering also cryogenic conditions. Membranes as PMD for storable propellants allow extended exposure time to fuel, similar improvement for oxidizers are in progress. Utilization of membranes that separate fluid and gas content of a tank are therefore promising, since the complex multi-phase flow image fields of free surface tanks, are avoided. The contractor shall assess, if present FSI methods, based on available material data of membranes, can reliably predict topology of membrane and kinematics of the fuel, forces on membrane and structures. Missing material data shall be identified and reported. Finally, a detailed analysis of the resulting margins of uncertainty for the coupled FSI result shall be provided. The contractor shall propose means to reduce the margins by improved modelling of liquid motion and / or structural response.
Background and previous activities Prediction of sloshing motion in tanks can be predicted well for large gravity. For capillary flows with small Bond numbers, surface tension and contact angles are presently not known already for ideal liquids, certainly not for cryogenic conditions. Numerically efficient solution methods, e.g. solving Navier Stokes equations for the kinematics of a fluid mass in a tank that is exposed to external acceleration require empirical data for contact angle and interface models that are presently not available for cryogenic liquids in micro gravity.
Technical Officer: Richard Schwane ([email protected]) 15 Toolset for the Design of Cryogenic Systems
Prog. TDE Budget 500K Duration: 24 months Aim: A Ref.: A 07 TRL: 2 to 3
Objective(s) Development and validation of tools for cryogenic systems, e.g. a cryogenic reservoir in orbit. Exiting tools, e.g. boil-off tools shall be critically reviewed, margins shall be identified and related to margins acceptable for design. Tools shall be updated by introduction of more accurate modelling methods and the result on the overall prediction error shall be estimated.
Description An existing and published integral tool, here for boil-off in a LH2 or methane reservoir in orbit shall be systematically checked for short-comings and updated to a level that allows application to complex cryogenic systems. Updates can include replacing empirical models, e.g. for heat transfer between liquid fuel and tank structures by more detailed CFD rebuilding, if justified and affordable. The dependency of prediction accuracy on time and particular nodal representation (e.g. number and arrangement of nodes) shall be quantified. Thorough validation of tool performance and accurate estimation of remaining margins of uncertainty is considered a pre-requisite for image successful tool application and is therefore considered the most important outcome of this activity.
Background and previous activities Presently, validated tools for design of complex cryogenic systems do not exist, but simplified tools, as the present boil-off tool are frequently used for optimization tasks. Tool application to design requires detailed knowledge on tool accuracy.
Technical Officer: Richard Schwane ([email protected]) 16 Material Point Method for Sloshing and Multiphase Flows
Prog. GSTP Budget 300K Duration: 18 months Aim: A Ref.: A 11 TRL: 4 to 5
Objective(s) Review, improve and document functionality, including accuracy, computational efficiency and robustness of dual Eulerian and Lagrangian CFD methods for the description of multi-phase flows.
Description: Presently, multi-phase flows are predicted with numerical schemes that solve for conservation of mass, momentum and energy with added empirical models for physical aspects of the solution that are not covered by conservation properties. In this activity conservation schemes (Eulerian description) and particle methods (Lagrange description) are combined in one numerical method that is not dependent on, often difficult to obtain, empirical models or data for missing physics. This activity shall review existing dual schemes, particularly MPM and SPH schemes for their image ability to rebuild complex multi-phase flows, propose and introduce scheme updates and demonstrate functionality as cost, accuracy and robustness of such advanced schemes, if compared to Navier Stokes and empirical modelling.
Background and previous activities Accurate description of even physically simple, i.e. isothermal multi-phase flows is still a challenge for conservation schemes, e.g. the Navier Stokes equations. Advanced numerical methods that include all physics required to rebuild the flow field and, simultaneously the free interface topology in the flow field without the need for additional empirical data from experiments are systematically examined for their potential to improve prediction accuracy/cost for such methods.
Technical Officer: Richard Schwane ([email protected]) 17 Improved Ullage Bubble Dynamics Model
Prog. GSTP Budget 300K Duration: 18 months Aim: A Ref.: A 19 TRL: 4 to 5 Objective(s) To improve physical models for ullage dynamics in micro or zero gravity for storable and cryogenic applications, to allow predictions of effects of ullage bubble relocation in the tank on spacecraft dynamics as input to a GNC system.
Description Prediction of kinematics of an ullage bubble in a tank in micro gravity is still a challenge. Since conservation equations for mass, momentum and energy do not contain information on molecular physics that are responsible for the, in micro gravity, dominant surface tension and contact angle effects on liquid and IF movement, empirical data has to be provided as boundary condition for the Navier Stokes equations. Particularly for cryogenic applications, such empirical data is today still not available in a generally usable and validated form. Resolution of all physical influences that determine bubble dynamics for cryogenic multi phase flows is still not at hand and will not be available anytime soon. A central part of this activity is therefore the quantification of expected solution accuracy for different methods with different complexity and costs, allowing to propose the most economical solutions methods that meets the accuracy image requirements, coming from design.
Background and previous activities Kinematics and dynamics of ullage bubbles is still not fully understood. Some progress has been made in recent years leading to the conclusion that dynamics of ullage bubbles in zero gravity cannot be predicted from linear dynamic models. Improved modelling is therefore required, selecting the most economical method that still considers physics sufficiently well to allow to predict the most characteristic features of sloshing in micro gravity and its effect on space craft dynamics.
Technical Officer: Richard Schwane ([email protected]) 18 Development of an Experiment for Transient Phase Change Aspects in Microgravity for Cryogenic media Prog.GSTP Budget 600K Duration:24 months Aim:B Ref.:B 02 TRL: 4 to 5
Objective(s)t This activity shall assess the status of experiments for transient phase change phenomena of cryogenic liquids in micro gravity. It shall identify and collect existing data and justify, plan, design and conduct experiments to supplement presently still missing data. Experiments shall be designed as breadboard, for existing micro gravity platforms, parabolic flight campaign or ISS, as needed and if justified. The contractor shall investigate in detail, in how far breadboard experiments can be used to obtain the required data for cryogenic liquids in micro gravity.
Description Phase change of cryogenic fluids in micro gravity, particularly for capillary flow conditions, is still not reliably predictable by available CFD methods. This is mainly due to unavailability of experimental data recorded under conditions that are sufficiently similar to flight. Such data is required to build and update physical models in CFD tools for phenomena that are not covered by the physics that is resolved by CFD method, e.g. molecular dynamics determining interface topology in multi-phase flow problems, which is not accounted for by mass, momentum and energy conservation as described by Navier Stokes equations. For cryogenic liquids, transient phase change phenomena for different time scales are responsible for many aspects of flow development. Due to the non-linearity image of the physical effects involved, even short duration transient phenomena can determine the flow development. Identification of such short term multi-phase phenomena that have a strong impact on the overall flow development, e.g. effects of the Leidenfrost effect on bubble attachment versus bubble rejection by a heated wall structure is of prime importance.
Background and previous activities Experimental data for phase change phenomena of cryogenic liquids in micro gravity is not available with an accuracy that would allow to use experimental data to build physical models for the consideration of such phenomena in CFD rebuilding.
Technical Officer: Richard Schwane ([email protected]) 19 Spacecraft Re-fuelling Ground Testing
Prog.GSTP Budget 1.2M Duration:24 months Aim:B Ref.:B 05 TRL: 4 to 5
Objective(s) This activity shall perform ground testing for spacecraft refueling. Testing shall focus on elements that can be justified to the sufficiently independent from the test environment, as mechanisms for docking and undocking, leak tight fuel line connection and disconnection with minimal risk of contamination of instruments on the receiving spacecraft.
Description Spacecraft refueling in LEO or GEO is presently considered a promising solution to allow for exploration missions beyond the moon. A reservoir for cryogenic fuels shall be placed in orbit, allowing spacecraft launched from the earth surface to replace fuel that has been consumed to escape earth gravity and extending their range of operation in orbit, substantially. Whereas some functionalities of a refueling device can only be tested in a relevant micro gravity environment, a number of tests can be performed in gravity, i.e. technology and operations required to dock spacecraft to reservoirs, to connect and test fuel lines and connectors for leaks, to disconnect fuel lines with minimal loss of fuel image to mitigate the risk of contamination of sensitive instruments, etc. The contractor shall identify testing that can be performed on ground and justify the test by thorough analysis of the governing similarity rules. For feasible tests he shall design a test facility, using available elements, sensors, robotics and mechanisms, as much as possible. Tests shall then be conducted, in line with the available budget. Recommendation for additionally required testing shall be provided.
Background and previous activities Spacecraft refueling has been routinely performed for storable propellants and consumables, e.g. to allow ISS extended operation in orbit. The present activity shall utilize the findings from such developments and assess, which additional requirements have to be met to extend fueling activities to cryogenic fuels and refueling of spacecraft from a LH2 or methane reservoir in orbit.
Technical Officer: Richard Schwane ([email protected]) 20 Instrumentation for Cryogenics Propellant Tanks
Prog.GSTP Budget1.5M Duration:24 months Aim:C Ref.:C 03 TRL: 4 to 5
Objectives: This activity shall assess the status of instrumentation for cryogenic tanks, select the best technical solution for each instrumentation option, justify the choice and recommend additional elements that might be needed for new applications.
Description Conventional cryogenic propellant tanks need to be instrumented to allow for fuel gauging and monitoring of tank functionalities, e.g. control of tank pressure and related boil-off. Instruments have to withstand the rough thermal conditions in a tank for the entire operational time, i.e. some years for fuel reservoirs in orbit. The contractor shall assess available instrumentation that is already verified for cryogenic conditions and identify additionally required hardware to meet the objectives. Special attention shall be paid to new functionalities, e.g. of tanks that can be re-fueled with cryogenic liquid after extended operation. Additional requirements and specifications for such hardware, e.g. for shall be identified and image reported. Optical fibers have been identified In breadboard experiments as a means to improve reliability and speed of signal transmission between sensors and control elements. This hardware shall be assess for its compatibility to flight conditions, including robustness to withstand the loads acting on sensors and optical fibers or electrical wiring.
Background and previous activities Instrumentation of cryogenic tanks with sensors and non-intrusive methods has been assessed in detail in the Cryosense activity /…/. Optical fiber based instruments have been validated for temperatures as low as 20K, some elements have been demonstrated to work in breadboard experiments in liquid Helium at 3K.
Technical Officer: Richard Schwane ([email protected]) 21 Ultrasonic Non-intrusive Measurement Techniques for Storable and Cryogenic Liquids Prog.GSTP Budget 300K Duration:18 months Aim: C Ref.:C 04 TRL: 4 to 5 Objective(s) To extend a recently developed non-intrusive measurement method based on ultrasonic sound fields to cryogenic conditions.
Description Non-intrusive measurement techniques are an important means to gain information on the condition of fuels in tanks during spacecraft operation. Particularly for cryogenic liquids penetration of the tank structure for cables or optical fibers can increase the complexity of a tank design, considerably and will increase undesired heat transfer into tanks. Accurate gauging allows efficient use of fuel during a mission. Non-intrusive measurement techniques, e.g. by means of transmission, reception and analysis of ultrasonic sound fields have been demonstrated to work under breadboard conditions for storable propellants. This development shall be extended, to allow for ultrasonic diagnostics also for cryogenic liquids. The contractor shall investigate in detail, if the technology can be applied also without tank penetration and on image already space qualified tanks. The measurement equipment that radiates, records and analyses the sound field shall be redesigned to allow for cryogenic temperatures. The contractor shall investigate, if non-penetrating mounding of transmitters and sensors is already feasible. If that is the case work shall be focused on developing the measurement equipment in terms of size, weight and implementation requirements to a level that allows to integrate it on existing space craft, If non-penetrating mounting is shown to deteriorate the quality of the measured data, focus of this activity shall be to improve the experimental hardware in this respect and to improve the design of cryogenic tanks to mitigate the effects of non- penetrating mounting.
Background and previous activities Non intrusive measurement methods based on transmission, recording and analysis of ultrasonic sound fields into Breadboard set-up from Cryosense a tank have been developed and demonstrated for breadboard applications and non-cryogenic conditions in /…/.
Technical Officer: Richard Schwane ([email protected]) 22 Consistent prediction of liquid/gas interfaces in micro gravity
Prog. GSTP Budget 500K Duration: 24 months Aim: A Ref.: New 01 TRL: 2 to 4
Objective(s) This activity shall investigate the limits of efficient numerical schemes, e.g. Navier Stokes equations for the prediction of capillary flow phenomena and propose strategies to mitigate the short comings to extend the validity to more and more capillary flow conditions.
Description Liquid/gas interfaces in micro gravity are determined by capillary mechanisms. Correct prediction of IF topology is a pre-requisite for CFD rebuilding of technically relevant flow phenomena, e.g. heating and related boil-off of cryogenic fuels in tanks and reservoirs, impact of sloshing on CoG, impact of slosh damping on pointing accuracy of exploration and telecommunication satellites, etc.. Whereas for high gravity and Bond numbers, empirical models that describe the pressure jump normal to an IF are sufficiently accurate to predict realistic interface dynamics, for small Bond numbers (Bo<1) flow predictions are affected by the inability of conservation schemes to consider criteria on IF topology coming from Lagrange image Potential, i.e. not allowing steady state solutions that do not guarantee that all particles are arranged, as required by minimization of the Lagrange Potential. This activity shall assess the severity of this short coming for selected applications. For critical applications, e.g. results from capillary testing in ISS for LIZARD, a concept shall be developed that can extend the validity of efficient CFD methods into the capillary flow domain. Results shall be validated against existing data from micro gravity testing, new experiments shall be proposed, where necessary and conducted in accordance with available budget.
Background and previous activities Prediction of capillary flow fields is required for design of a large number of currently planned new applications, e.g. design of fuel reservoirs in orbit and related refueling of space craft in micro gravity. For capillary flow conditions, efficient numerical methods have failed to predict even isothermal flow phenomena. Activities are ongoing to improve the quality of required empirical data, e.g. for IF dynamics and contact angles. Utilization of such improved empirical data requires Navier Stokes solvers with quantifiable prediction errors.
Technical Officer: Richard Schwane ([email protected]) 23 Uncertainty quantification for cryogenic flow modelling in micro gravity. Prog. TDE Budget 400K Duration: 18 months Aim: A Ref.: New 02 TRL: 2 to 4
Objective(s) Quantification of approximation error for CFD predictions for cryogenic flow problems in micro gravity including effects of required empirical models.
Flow rebuilding with CFD methods contains an approximation error that depends on chosen discretization of the spatial and temporal scales of the problem. If solutions are entirely determined by conservation principles, e.g. for mass momentum and energy as for the Navier Stokes equations, reliable methods are available to characterize the prediction error and reduce it, if required by optimization of the prediction set-up to any desired value. In case the solution is significantly influenced by empirical models, e.g. for flow features that are not covered by conservation principles, the total prediction error also depends on the quality of empirical data. For such more complex systems, a method shall be developed that allows to quantify overall prediction errors, including effects of empirical modelling or unavoidable model inconsistencies of existing CFD methods. Based on such findings that contractor shall provide means that allow to select simulation parameters as spatial and temporal resolution and, e.g. required accuracy of empirical input data to guarantee a pre-defined error margin on data of technical interest.
Background and previous activities CFD methods for multi-phase applications, even for isothermal and incompressible liquid/gas problems, are still suffering from not well understood margins of accuracy, due to utilization of empirical models. For cryogenic applications margins of experimentally recorded empirical data are generally enlarged and can accumulate. Particularly for such applications, error margins for results of technical interest should be available before start of the simulation.
Technical Officer: Richard Schwane ([email protected]) 24 CFD and breadboard testing for empirical data for non- isothermal applications Prog. TDE Budget 300K Duration:18 months Aim: A Ref.: New 03 TRL: 2 to 3
Objective(s) To extend an existing breadboard test facility to include investigation of non-isothermal tow phase flow problems.
Description In this activity the domain of application of a breadboard experiment that allows to record relevant flow data for applications in micro gravity shall be extended to non-isothermal single phase flow problems. Existing hardware shall be adapted, allowing, e.g. to accurately prescribe surface temperatures independent from the temperature of the liquid. Temperatures shall be electronically controlled, provisions to keep the liquid temperatures constant over time, shall be investigated, if necessary. A CFD method shall be selected that has been, at least, partially validated for the prediction of heat transfer in image incompressible two phase flows. The method shall be validated in detail for accuracy of heat transfer predictions. The validated method shall then be utilized to predict experimental results using existing empirical models for static and dynamic contact angles and hysteresis, the deviation of measured and predicted flow parameters shall then be used to optimize and fine-tune empirical CA models using available System Identification (SI) methods and tools.
Background and previous activities A breadboard test facility has been assembled at ESTEC that allows to quantify the effect of physical mechanisms that are normally only visible in micro gravity also in ground experiments. Proof of concept has been demonstrated for isothermal flows, however, the method is as well applicable to multi-physics application, e.g. including non- isothermal effects on the flow field.
Technical Officer: Richard Schwane ([email protected]) 25 Acoustic manipulation methods for cryogenic flows
Prog. TDE Budget 600k Duration:24 months Aim: B Ref.:New 04 TRL 2 to 3
Objective(s) This activity shall extend the range of application for a novel technique that can move fluid and gas by means of ultrasonic sound fields from an array of piezo electric transducers. It shall check and confirm the robustness of hardware in cryogenic conditions and demonstrate this promising flow control method for cryogenic applications, in a breadboard environment.
Description Based on available promising results for isothermal applications, the technique shall be extended to cryogenic conditions for application to existing cryogenic launcher U/S, orbital reservoirs and refueling of spacecraft in orbit. For cryogenic application, heat intake into the tank requires complex and expensive cooling or boil-off of fuel Therefore the method has to be carefully designed to limit heat intake to a level that can be justified by the benefit of the fluid manipulation method. image: (current
Background and previous activities Acoustic manipulation techniques have been demonstrated to efficiently move isothermal liquid and ullage bubbles in bread board experiments, mitigating potential risks of bubble ingestion in fuel lines to the engine. Based on available promising results for isothermal applications, the technique shall be extended to cryogenic conditions for application to existing cryogenic launcher U/S, orbital reservoirs and refueling of spacecraft in orbit.
Technical Officer: Richard Schwane ([email protected]) 26 Cryogenic aspects of refuelling
Prog.TDE Budget 500K Duration:24 months Aim: A Ref.:New 05 TRL: 2 to 3 Objective(s) This activity shall extend the technologies developed for refueling of storable propellants in orbit to cryogenic liquids.
Description Assessment of presently available engine technology in the context of Tsiolkovsky’s rocket equation indicates that fuel reservoirs in orbit are required for exploration missions beyond the moon. Presently ongoing activities on prototyping of refueling with storable propellants in LEO and GEO shall be extended to allow operation with cryogenic liquids. Cryogenic liquids can undergo phase transitions during refueling, that require special attention since requirements, e.g. for gas free expulsion of liquids from a refueled tank, and venting of tanks during re-filling still have to be met. This activity shall investigate, which tools are already available and validated to support the design. Existing shortcomings shall be identified and related updates shall be proposed and implemented, if consummate with the available budget. The contractor is expected to provide detailed recommendations on additionally required development steps and require future activities.
Background and previous activities Prototyping of refueling in orbit is ongoing, focusing on applications for storable propellants. New applications are emerging, e.g. allowing extended exploration missions by refueling of space craft in LEO or GEO.
Technical Officer: Richard Schwane ([email protected]) 27 Development of a Non Intrusive Sloshing and Refueling Spacecraft Demonstrator for Micro-gravity
Prog. TDE Budget 2000K Duration:24 months Aim: D Ref.: D 02 TRL: 2 to 4
Objective(s) To develop, design and validate a demonstrator for sloshing and refueling in micro gravity with non-intrusive measurement technology.
Acoustic manipulation techniques have been demonstrated in a breadboard environment for isothermal flow problems. The technique shall be extended to cryogenic onditions and be demonstrated as a means for, e.g. bubble removal from cryogenic liquids in micro gravity.
Background and previous activitiesNon-intrusive measurement technology has been developed and validated for storable propellants in a bread board environment. Currently this activity is extended to allow for flow field diagnostics also for cryogenic fuels. The present development shall be demonstrated on a demonstrator flight using sounding rockets.
Technical Officer: Richard Schwane ([email protected]) 28 Appendix: Cryogenics in Microgravity
Outline – Needs, status and perspective for future activities • Brief Assessment of new missions and requirements • What is now needed ? • What is already working ? • What is presently still missing ?
29 Appendix: Cryogenics in Microgravity
Example for new requirement: Fuel reservoir in orbit. • Large dimensions • Capillary conditions • Cryogenic fuels ?
Experiments: • Full scale flight testing is expected to be cost-prohibitive • Ground micro g facilities have limited test duration • Breadboard activities could allow specific micro gravity data at reduced costs
https://www.energy.gov/sites/prod/files/2014/03/f9/compressed_hydrogen2011_11_chato.pdf Simplified tools, e.g. boil-off tools: • Not completely validated: Cross validation between simplified models missing • Relevant physics is partially excluded to maintain validity of simplified modelling (mixer is assumed to avoid stratification) • Required modal resolution (number of liquid and structural nodes) for tools is not clear: …
30 Appendix: Cryogenics in Microgravity
CFD methods: • DNS/DSMC: working and validated but too expensive • Navier Stokes: requires additionally empirical models for physics that is not resolved by conservation schemes for isothermal, non-isothermal and non-isothermal with phase transitions : • surface tension (ST), • static and dynamic contact angle (CA) and hysteresis • Consistency issues exist with ST modelling for low Bond numbers • Dual Eulerian and Lagrangian methods (e.g. MPM and specifically SPH) do not need additional measurements for empirical modelling of ST and CA, but resulting solution accuracy has not yet demonstrated clear improvement over Navier Stokes solutions.
31 Appendix: Cryogenics in Microgravity
• Conclusions:
• Simplified methods (boil-off models) not mature for design
• Experiments in relevant similarity conditions, particularly requested for empirical data for conservation schemes, i.e. Navier Stokes, are not expected anytime soon
• CFD suffers from missing empirical data and inconsistent boundary conditions, e.g. for ST in capillary flows
To do:
• Simplified methods (e.g. boil-off tools) have to be matured and validated for complex applications. Methods should be validated as far as possible, margins of uncertainty have to be quantified and reduced, by replacing over-simplified modelling by more advanced methods, e.g. CFD tools.
• Breadboard testing should be developed to allow recording of required empirical data with approximated micro gravity conditions, where possible. Strict requirements shall apply.
Stronger focus on utilization of breadboard testing, as long as margins of uncertainty are understood and CFD is available and validated (grid and time converged) for the selected set-up
+ Such experiments could be easily extended to multi-physics, e.g. for CA testing in non-isothermal conditions
32 Appendix: Cryogenics in Microgravity
• CFD has to be matured, boundary conditions, e.g. for CA and ST used in multi-phase modelling has to be checked for consistency • Remaining shortcomings (e.g. for ST formulation) shall be mitigated, as far as possible • Flow phenomena that presently cannot be resolved and predicted with Navier Stokes shall be modelled, e.g. : • Onset of boiling for non-isothermal liquids in tanks, • Attachment and detachment of ullage gas and liquid lumps to tank walls. • …
33 Development and Flight Validation of Multi-physic Tools for Propulsion Controlled Landing Prog.GSTP Budget 500 Duration:24 months Aim: D Ref.: D 03 TRL: (current to Target) Objective(s) Objective of the present activity is the development and validation of an inter-disciplinary prediction tool that can take all relevant physical mechanisms during propulsion controlled landing in earth of planetary atmosphere into account. After validation of the single discipline components of the methodology the tools shall be coupled in a multi-disciplinary prediction tool that will allow to optimize the design of the hardware for robustness and stability. A subscale lander shall be assembled from off-the-shelf components. Application of the multi-physical tool to the subscale demonstrator allows to validate the multi-disciplinary tool and demonstrate the robustness of the algorithms in flight. This activity will also comprise to input of experimental data"
Description This activity shall assess the status of multi-physics predictions for descend and landing on planets (including earth), moon and asteroids. Depending on spacecraft mass, gravity level and atmosphere, propulsion controlled landing can be required. This activity shall develop and validate a multi-physics tool that can keep track of, at least, the most important physical effects that have an influence on descend and landing of a spacecraft.
Background and previous activities Propulsion controlled landing can be required for exploration missions.
Technical Officer: Richard Schwane ([email protected]) 34 Development of a Flight-back Booster Demonstrator Phase B
Prog. GSTP Budget 3000K Duration:36 months Aim: D Ref.:D 06 TRL: 3 to 5
Objective(s) Development of a Flight-back Booster Demonstrator
Description Critical assessment of State-of-the-Art of flight-back-booster technology. Identification of technology gap, risk assessment.
Background and previous activities k Continuation off Flight back booster demonstrator development.
The LFBB model used in wind tunnel tests by the German Technical Officer: Richard Schwane ([email protected]) Aerospace Center (DLR) 35 Green Propellant Usage in Long Duration Exploration Missions TDE 400 k€ Duration: 24 months Aim: A Ref.: - TRL: 3 to 4
Objective(s) The reduction of stability and storability, because of (adiabatic) decomposition or degradation due to surface reactions, of low- toxicity/green propellants needs to be characterised in order to use them safely and efficiently in long-duration exploration missions, in particular crewed missions. The current basic research activity will improve the understanding of the material compatibility and storability of green propellants, which will facilitate the selection of propellant and material combinations during mission and spacecraft design.
Description In the last few years there has been a push to develop low-toxicity, so-called green, propellants (e.g. hydrogen peroxide, ADN-blends, etc) as replacements for standard hydrazine-based systems. This is in part because of Europe's "Registration Evaluation Authorisation and Restriction of Chemicals" (REACH), which has put hydrazine on its list of high-concern substances, but also because of the reduced environmental impact and potentially lower operational cost (due to less health & safety restrictions). In particular for long-duration missions that include a potential human presence, such as future missions to the Lunar Gateway, the usage of low-toxic propellants could be highly beneficial. Long-duration exploration missions require highly storable and stable propellants that can remain safe and effective for several years. The storability and stability of low-toxicity propellants, such as hydrogen peroxide, is driven primarily by surface decomposition. Experimental testing of these characteristics should therefore be done for combinations of propellants and surface materials, to allow for the appropriate selection of tank and/or liner materials to be used with those propellants. Such testing can give rates of propellant degradation/decomposition (or Active Oxygen Loss in the case of hydrogen peroxide), but high-quality testing data is scarce. © European Astrotech Ltd The current activity will include an extensive review of low-toxicity (green) propellants, to identify the driving parameters, risks and constraints for the selection of such propellants for long-duration exploration missions. In particular, data on propellant ageing/degradation and (adiabiatic) decomposition will be gathered, and dataset gaps identified. An experimental campaign to test the decomposition and material compatibility of low-toxicity propellants will be designed and performed, which will allow a more accurate characterisation of the different propellants and their suitability for long-duration exploration missions.
Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (A22)
Technical Officer: Jeroen Van den Eynde ([email protected]) 36 Post Flight Data Analysis Toolset
GSTP 400 k€ Duration: 12 months Aim: A Ref.: - TRL: 4 to 6
Objective(s) The objective of the activity is to develop a software tool for level-2 post flight data analysis to support flight reconstruction and confirmation of flying qualities.
Description The post-flight data analysis of a flight vehicle is an important element in the assessment of its nominal and/or off-nominal performance and behaviour. Different levels of post-flight analysis can be performed, depending on the level of detail required and disciplines involved. A toolset for level-1 post-flight analysis is currently under development, to be released publically as an open- source code, which deals with the disciplines of propulsion, aerothermodynamics, performance and structures.
The current activity does not envisage to provide an update of this toolset, but rather focus on the development of a more in-depth and detailed level-2 analysis methodology of the flight vehicle performance. Interfacing the two toolsets could however be foreseen in the activity.
The level-2 post flight data analysis toolset will be able to provide a detailed flight reconstruction (i.e. obtain the Best Estimated Trajectory), system identification (i.e. model reconstruction of the dynamic and kinematic behaviour) and confirmation of the flying qualities (i.e. obtain an accurate mathematical description of the behaviour and vehicle performance).
The activity will entail the design and development of the level-2 analysis methodology, and its validation using flight-data.
Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (A27)
A level-1 toolset for post-flight analysis (PFAT) is currently being developed under the ESA TDE programme.
Technical Officer: Jeroen Van den Eynde ([email protected]) 37 Gas-Surface Interaction and Soil Erosion due to Pulsed Jets GSTP 500 k€ Duration: 18 months Aim: B Ref.: - TRL: 3 to 5
Objective(s) The goal of the activity is to understand the dynamic gas-surface interaction of pulsed jets impinging on soil underground, characterise the soil erosion and uplifting, and identifiy their effects on a spacecraft during terminal landing.
Description With the new interest into the Lunar Gateway and the potential of recurrent ascent/descent operations on the Moon’s surface by European spacecraft, e.g. European Large Logistics Lander (EL3), an improved understanding of the plume/surface interaction with granular flows is required. In particular the dynamics of this interaction, the pressure oscillations of the plume flow and resulting soil behaviour needs to be understood. While it has previously been shown that some elements of the plume/surface interaction can be simulated using porous media equivalent models (i.e. Darcy’s Law), the erosion and uplifting of the soil requires the modelling of the granular underground.
The envisaged activity would entail the high-fidelity modelling and simulation of the gas/surface interaction with the inclusion of the soil (erosion, cratering, uplifting) under the influence of pulsed jets in a variety of relevant spacecraft conditions. Small-scale experimental tests are foreseen to validate the modelling and simulation accuracy. The results of the simulations would be used to characterise the granular flow dynamics over short timescales and the potential feedback effects on the spacecraft during landing. A low-order empirical model of this extensive characterisation will be derived to facilitate the early design phases of planetary landers.
The characterisation activity will eventually aim to answer the following questions, among others: 1. What is the size distribution, volume and velocities of soil particles lofted by exhaust plumes under different conditions and configurations? 2. How does the landing zone surface change due to the plume/surface interaction due to recurrent landing operations?
Background and previous activities The resulting simulations will enable future cross-verification and validation activities using an experimental test facility currently under development. This facility will allow the testing of plume/regolith interactions under vacuum conditions of Mars and airless bodies (e.g. © CFD Research/NASA MSFC/Jacobs Space Exploration Group Moon).
Technical Officer: Jeroen Van den Eynde ([email protected]) 38 Hypersonic Shockwave/Boundary-layer Interaction on Ramps and Flap Deflections TDE 300 k€ Duration: 12 months Aim: B Ref.: - TRL: 2 to 4
Objective(s) The objective of the activity is to characterise shockwave/boundary-layer interaction, and in particular the behaviour of the recirculation region, at hypersonic speeds on a variety of geometries, representative of flap deflections and/or other geometric features of flight vehicles.
Description Control surfaces and geometric features on hypersonic flight vehicles induce strong shockwave/boundary-layer interactions (SBLI) that could directly affect the required design due to thermal and structural loads, and could impact the controllability and flying qualities of the vehicle. In particular shock-induced separation and reattachment could produce energetically-significant low-frequency fluctuations and increased structural heating, which could potentially lead to undesirable loads on the structures and decreased flap efficiency.
While a large number studies have previously addressed the topic of SBLI on compression ramps, limited research has been performed on the effect of a detached compression ramp, i.e. representing flap deflection with a hinge gap. Such a hinge gap could be © ESA/ASI/RUAG/Thales present on flaps of real high-speed flight vehicles, and the effects should therefore be understood.
The current activity would look into the dynamics of hypersonic shockwave/boundary-layer interaction of ramps/flaps, considering the effects of the gap size, deflection angle and incoming boundary-layer state (laminar, turbulent, transitional).
Experimental wind tunnel tests and high-fidelity numerical simulations are expected to be performed, which will yield an extensive characterisation of the shockwave/boundary-layer interaction behaviour.
Background and previous activities Previous activities have looked into SBLI on flat plates and compression ramps, but the effects of hinge gaps on the recirculation region and its dynamics has not been extensively studied.
Technical Officer: Jeroen Van den Eynde ([email protected]) Priebe & Martin (2012), JFM Vol. 699, pp. 1–49 39 Multi-metal additives in solid rocket motors
FLPP 300 k€ Duration: 12 months Aim: B Ref.: - TRL: 4 to 5
Objective(s) To investigate the effect of solid propellant multi-metal additives (i.e. Al-Mg) on a rocket motor performance, and the potential impact of the resulting plume on the atmospheric environment.
Description It is very well known that metal additives (e.g. aluminium, boron, etc.) in solid rocket motor propellants can improve the combustion, performance and/or characteristics of the motor and plume. However, the addition of aluminium, often used in HTPB/AP-based solid propellant, and combustion, results in aluminium oxide (alumina) in the plume. Alumina particles can have a significant effect on stratospheric ozone depletion by initiating chlorine activation reactions. Saile et al. (2019), FAR Conference
It has previously been observed that some metal additives can reduce the plume chlorine content significantly, and thereby possibly counteracting the ozone depletion potential of the plume, decreasing the potential adverse effects on the atmosphere and environment. For example, the addition of magnesium particles to the propellant mixture can enhance the chlorine reduction, but might have a negative effect on the combustion and/or performance. Combinations of metal additives (e.g. Al-Mg), either as mixed or coated particles, could therefore be found to potentially benefit both the performance and environmental effects.
The envisaged activity would entail the testing of multi-metal additives on solid propellant performance, and characterisation of the resulting plume contents. The potential effects of the plume content, and therefore of the multi-metal additives, on the ozone depletion potential in the atmosphere will be characterised.
Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (B01)
Previous ESA-funded work has looked into the alumina content of solid rocket motors for various propellant formulations [Experimental Modelling of with Alumina Particulate (EMAP) in Solid Rocket Motors – DLR/PoliMI/FOI].
Zhen et al. (2019), RSC Adv. Issue 33
Technical Officer: Jeroen Van den Eynde ([email protected]) 40 0D and 1D Aerothermodynamics Tools for Accurate Modelling GSTP 300 k€ Duration: 12 months Aim: A Ref.: 02 TRL: 4 to 5
Objectives: To develop 0D/1D models that can be in conceptual design, used as stand-alone elements or linked to other engineering tools. To link these with an accurate computational fluid dynamics modelling software.
Description From conceptual design to operation we need to employ lower-order aerothermodynamics models for different applications; in conceptual design most of the system model is composed of low-order or re-used elements from earlier studies. During detail design and operations, the focus shifts to predicting dynamic behaviour of the element/system at different operational and off-nominal conditions. This activity has two goals; to expand the current array of 0D/1D modelling tools to cover more of the engineering-level analysis needs (mainly for early design phases) and secondly, to link these tools to high-fidelity analysis tools, where they provide the boundary conditions.
Modelling areas identified: • Engineering-level methods for predicting forces and moments on launch vehicles, re-entry vehicles and suborbital hypersonic aircraft using empirical and semi-empirical models (for example, Drag Breakdown Methods) Image source: Detroit Engineered Products • TPS shock, ablation and heat flux models • Acoustics models to predict environment under launcher fairings • Ground-plume interaction to model parasitic forces and moments on landers and ascent vehicles, ground pressure (drives soil erosion and ejected dust-related phenomena). • Models for forces, moments and heat fluxes for retro-propulsive maneouvres of launch vehicle stages Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (A02). The elements developed in this activity should target compatibility with existing industrial system modelling and optimisation environments and tools (e.g. ESPSS, ASTOS, Simulink, Modelica). A candidate implementation would be through the Functional Mock-up Interface (FMI) standard, which describes a standardised container and interface to exchange dynamic models and is already supported by many of these software. Image source: Empresarios Agrupados/DLR
Technical Officer: Csaba Jéger ([email protected]) 41 Fully Automatic Volume Mesh Generation for Fluid Dynamics TDE 300 k€ Duration: 18 months Aim: B Ref.: - TRL: 1 to 3 Objective: To demonstrate the generation of high quality unstructured volume meshes for arbitrary geometries and boundary conditions using machine learning methods. Description Automatic mesh generation has been long sought-after in Computational Fluid Dynamics (CFD). The generation of body- fitted volume meshes required by most schemes is still a labour-intensive, manual process. Besides post-processing, mesh generation is close second in man-hour requirements. CFD solutions are heavily influenced by the structure and topology of the numerical mesh and constructing an suitable one for a given problem requires certain a priori knowledge about the structure of the flow, which is not feasible to automate using classical algorithms for complex real-world geometries. Machine learning (ML) techniques could overcome this limitation and replace human interaction entirely for numerical mesh generation. It has already been demonstrated that these methods are able to roughly predict flowfields for sub-sonic cases purely based on the geometry and boundary conditions for negligible computational effort.
This activity aims to investigate the application of machine learning methods for general flow computations, focusing on flowfields with mixed subsonic and supersonic regions in the continuum flow regime. Machine learning techniques are applied to predict flow features requiring refinement which is then used as an input to size the volume mesh.
Background and previous activities Flow features in mixed subsonic-supersonic problems are diverse and include rarefaction-recompression waves, recirculation, flow detachment, shockwave impingement and many others. A successful method will need to learn to extract these general features and develop an internal approximation of them to be successful on arbitrary geometries and flow conditions.
An important part of mesh generation is the sizing of the prismatic boundary layer mesh where a sufficiently small first-cell height is required for accurate boundary layer resolution, but the exact size is the function of local flow conditions. Correct sizing of this region will be one of the key performance metrics during evaluation. Performance evaluation is quite straightforward as the CFD solution is considered as ground-truth so we do not deal with the problems of going from the synthetic training set to real-world data.
Technical Officer: Csaba Jéger ([email protected]) Image source: https://doi.org/10.1145/3197517.3201325 42 Multi-disciplinary Tool Coupling for Flight Vehicle Engineering GSTP 400 k€ Duration: 18 months Aim: A Ref.: 04 TRL: 3 to 4
Objectives: To perform multi-disciplinary breadboard modelling of a flight vehicle, including all major subsystems (propulsion, thermal, structural, avionics, power, life support, communications, trajectory, aerodynamics) using a coupling of existing tools and models. Description This activity aims to improve interoperability of currently separate models for all major vehicle subsystems. There has been tighter coupling and existing solutions to couple flow simulations with structural and thermal simulations (Fluid-Structure Interaction, conjugate heat transfer problems) with mature software tools, but further interaction with other subsystems is mostly done through ICDs. In upcoming mission and vehicle concepts, it is desirable to achieve earlier and deeper interaction between these models.
The activity would build out the interfaces and software to couple different domain-specific solvers and models to produce a full, detail model of a flight vehicle with the aim to demonstrate end-to-end mission simulation. The target benchmark case is envisaged to be a conceptual launch or re-entry vehicle, or a powered lander (manned or unmanned). The overall design can be derived from existing examples. Of particular interest is the full simulation of the propulsion system, including power and hydraulic subsystems, sensors and mechanical actuators and their influence on vehicle external environment, structural loads and thermal environment. The overall fidelity and granularity of the end-to-end demonstration will depend on the performance of individual elements, with probably the flow solver being the critical element.
The activity should leverage existing interconnections between solvers and develop an interface where new components can be brought in with minimal overhead. A candidate implementation would be again through the Functional Mock-up Interface (FMI) standard. This would also enable synergies with the 0D and 1D Aerothermodynamics Tools activity.
Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (A03).
Technical Officer: Csaba Jéger ([email protected]) 43 Validation of methods for rarefied-flow aero- thermodynamics TDE 500 k€ Duration: 12 months Aim: D Ref.: - TRL: 3 to 4 Objectives: To measure, under representative conditions in a wind tunnel, the aerodynamic forces, moments and heat flux experienced during low-altitude operations (aerobraking, drag-free satellites) phase of several satellite models and planetary atmospheres. To use the test results to validate the current model of the design and engineering tools.
Description The aerobraking phase is a major element which enabled various ESA missions, namely Venus Express and ExoMars 2016 and played a major role in the GOCE mission. This technique is employed in e.g. the prospective EnVision mission with the aim to lower the overall mass of the spacecraft.
The aero-thermodynamic performance of spacecraft manoeuvres in rarefied hypersonic flow in the vicinity of planetary atmospheres has been studied. Nonetheless, the behaviour of the complex spacecraft geometries considered in real- wold missions have not yet been well studied experimentally, the results relying on simulation tools with low fidelity models.
Image source: MARHy/ICARE The activity shall bridge the identified uncertainty in analysis capability for the future missions by conducting an experimental campaign in a rarefied flow facility. The data to be obtained shall be the aerodynamic forces, moments and heat flux on a representative spacecraft and at relevant flight conditions to cover future mission needs, including VLEO Earth observation satellites, aerobraking missions at Mars, Venus and the gas giants. The results from the test campaign shall be used to validate high-fidelity numerical tools, which so far have only been validated for simple shapes.
Background and previous activities Due to the limited number of test facilities available, the representative satellite geometry and flow conditions have to be tailored to existing capabilties. Depending on available instrumentation, additional flow features could be of high interest, especially flow composition changes close to the vehicle surface with respect to freestream conditions. A need has been identified to be able to perform precision composition measurements at low altitudes.
Image source: DOI 10.13009/EUCASS2019-775 Technical Officer: Csaba Jéger ([email protected]) 44 Tools for the design, analysis, and validation of ablation and erosion processes during heat transfer Prog. TDE Budget: 500 K€ Duration: 12 months Aim: A Ref.: 25 TRL: 1 to 3 Objective(s) To develop a toolset for the prediction of ablation and erosion processes of a spacecraft Description Ablation and erosion are first order effects in the design of aerospace systems. In this activity, we propose to develop, implement and validate numerical tools to model and predict the ablation phenomena occurring during a destructive re-entry. Currently, ablation and erosion models for destructive re-entries remain unreliable compared to models used for the thermal protection systems or internal flows. The main weaknesses come from the low fidelity resolution of the flow and lack of proper modelling of the ablating materials. For now, most numerical ablation tools use simple melting criteria to model ablation while it has been demonstrated in several experimental studies that it is not the case. This activity will develop several tools to model the flow heat flux that will in turn be used to predict heat transfer, material changes and ablation. The tools will either be integrated into a high fidelity predictive re- entry tools or sued as a validation tool for low fidelity simulations. The validation of the tools will be performed using plasma wind tunnel facilities.
The following work logic is proposed: Task 1: Review of the driving ablation and erosion phenomena involved in destructive re-entries including spillage, melting, oxidation, spallation, and pyrolysis outgassing or typical classes of materials found in re-entering spacecraft (CFRP, GFRP, metals, glue etc.). Review of existing ablation tools for TPS. At this stage, the level of fidelity required to accurately predict demise shall be assessed. A critical review of existing numerical models shall be performed and the technical requirements for the tools (level of accuracy, modelling gaps, computational load) will be derived. A review of experimental methodologies to validate ablation models. Task 2: Design and implementation of numerical tools based on the requirements derived from the literature review. Major classes of materials such as oxidizing metals and composite materials at least shall be considered. The tool output shall be the effective mass loss, phase change and chemical alteration of the material along the trajectory. In parallel to the tool design, the need for further material characterization shall be assessed. The implementation shall rely, whenever possible on existing flow simulation tools and existing ablation tools. Task 3: Plasma wind tunnel testing and material calibration. In this task, material testing will be performed to calibrate the new tools developed. Finally, the toolset will be validated at material and equipment level following the recommendation described in the DIVE reference.
Technical Officer: Orr Cohen ([email protected]) 45 Derisk Assessment of an Air-Breathing-Electric Propulsion (ABEP) system for a CubeSat prospecting Mars Prog. GSTP Budget: 600 K€ Duration: 24 months Aim: B Ref.: New TRL: 3 to 4 Objective(s) Utilization of Air-Breathing-Electric Propulsion system for a remote-sensing cubsat on Mars, for prospecting its surface. Description The cubsat design will include the utilisation of an innovative Air-Breathing-Electric Propulsion (ABEP) technology to maintain the orbital elements of its orbit. The Mars atmosphere is primarily composed of CO2 (95.32%), as well as slight presence of nitrogen and argon. The use of ABEP technology for an atmospheric flight on Mars would use the residual atmosphere and allow the removal of on-board propellant as this propulsion system would use the Mars atmosphere as a propellant. The ABEP ingests the gas molecules from the Mars atmosphere, then ionizes and accelerates the ions to high exhaust velocities. The activity will focus specially on flight vehicle and aerotherodynamics environment of cubesat and the miniaturisation technology required for the ABEP propulsion system. NASA’s MABHET Concept The following work logic is proposed: (source: Task 1: Mission assessment and configuration design. https://www.nasa.gov/pdf/636899main_H Task 2: ABEP prototype design and miniaturization for cubsat. ohman_Presentation.pdf) Task 3: ABEP ground test planning. Task 4: On-ground testing
Technical Officer: Orr Cohen ([email protected]) 46 Assessment of a very-high cargo Mars lander vehicle
Prog. TDE Budget: 200 K€ Duration: 12 months Aim: D Ref.: New TRL: 1 to 3 Objective(s) To assess the system design with an emphasis on the EDL system of a very high cardo vehicle to land on the Martian surface Description With the recent increase of interest in high cargo space transportation within the solar system, multitude of exploration missions studies are planned or undertaken with the aim to further human knowledge of the solar system, as well as establish human presence, in particular on Mars. In previous activity an assessment of a Very High Power Cargo Transportation System to Mars was performed. The current activity shall cover the lander vehicle concept of the heavy cargo from Mars orbit to the Martian surface, departing from the aforementioned transpiration vehicle. This shall include the mission concept, the development of an EDL strategy, preliminary design of the EDL system and its feasibility. In addition, the assessment of the following subsystems will be performed: GNC and data handling, power, with a main focus on the propulsion, aerodynamics and aerothermodynamics. Furthermore, a preliminary configuration will be proposed, including trade-off of the different subsystems.
NTR-propelled Crew transport vehicle Background and previous activities (CTV) Copernicus in LEO – Mars DRA In a previous study ‘Assessment of a Very High Power Cargo Transportation System to Mars’ performed by University (Borowski et al., 2012; Drake, 2009) of Strathclyde, together with the University of Birmingham and the University of Southampton an assessment was made about such transportation system to Mars. The work focused on a preliminary flight vehicle engineering model design of the transportation system, including the vehicle configuration, and the development of preliminary simplistic numerical models. The models covered the scaling of the mass, propulsion (nuclear engine), habitat structure and consumables, as well as structural analysis of the truss. In addition, the overall mission performance analysis and optimization was undertaken.
Technical Officer: Orr Cohen ([email protected]) 47 New generation of sensors and instruments for aerothermodynamics Prog. TDE Budget: 400 KE Duration: 18 months Aim: C Ref.: 5152 (STAT) TRL: 1 to 3
Objective(s) This activity will research the current state of the art on sensors and measurements techniques and it will propose new ones. The aim is to prepare for the development of a new generation of sensors and measurements techniques developing instrumentation that is smaller, has less mass and volume and it is able to measure faster, quicker, and cheaper than the current generation.
Description The sensors and measurement techniques shall cover the following: pressure, heat fluxes, heat load, chemical composition, boundary layer state / turbulent transition, thermal protection erosion, atmospheric density, Mach number, surface-chemistry, thermal protection ablation, in-depth material response, and spacecraft attitude TPD recession sensors
The tasks of the activity shall be: Task 1: Analysis. This task will research the current state of the art on sensors and measurements techniques and it will propose new ones. After an initial research of the current state of the art, a trade-off among the new proposed sensors will be performed weighting costs and benefits so to select few concepts to be further elaborated. Radiative Heat Flux Task 2: Design and preliminary development of a selected set of sensors. In agreement with ESA, the task will design, preliminary development, and functional verification of a prototype selection of sensors. The selected sensors/measurement techniques will then be designed and preliminary tested to verify the functioning, confirm the level of improvement reached within the activity and propose possible further upgrades.
Task 3: Testing a prototype selection of sensors in a functional representative scenario. Evaluation of test results and creation of a development roadmap for technology maturation.
Background and previous activities The concept and approach to technologies is applicable to any exploration mission where aerothermodynamics measurements are needed or would be beneficial.
Pressure Technical Officer: Luca Ferracina ([email protected]) Thermopile sensors 48 Pre-development of a Miniaturised Advanced Common Atmospheric Probe (MACAP) Prog. GSTP Budget: 800 KE Duration: 24 moths Aim: D Ref.: New TRL: 4 to 6
Objective(s) Pre-development of a standard atmospheric entry probe with a common core part that can be adapted to different mission destination / profiles.
Description Pre-development of a standard atmospheric entry probe that can be used for landing in several planets and moons. The probe should be sized between 6U and 12U CubeSat. The mass should be less than 80 Kg. The MACAP shall be designed as a secondary payload concept for future missions to possible different planets with the primary scientific objectives to perform in-situ measurements of the planet atmosphere during the descent.
The tasks of the activity shall be: Task 1: State of the art . This task will research the current state of the art to clarify the need for this product with similar design while been sufficient flexible to allow the investigation of different planets. Task 2: Conceptual design. This task will focus on the identification and characterization of possible MACAP concepts looking at the major systems (i.e. Payload / instrumentation, Thermal Protection System, Communication) and their conceptual design Task 3: Preliminary requirements definition Task 4: Pre-development: This task will be dedicated to the design consolidation, preliminary development, and functional verification of a prototype of the selected MAPAC concept. The selected solution will then be preliminary tested to verify the functioning and propose possible further upgrades. Evaluation of test results and creation of a development roadmap for technology maturation.
Background and previous activities There are different planetary atmospheres (e.g. Earth Mars, Venus titan, Jupiter and Saturn) in the solar system exhibiting a wide range of behaviors. Entry probes working in coordination with orbital module could play an important role on their study. This is an opportunity to re-evaluate how to study (and observe) planetary atmospheres. Multible probes could be flown in the same entry mission to sample the atmosphere at different position at the “same time” International Planetary Probe Workshop – IPPW 10 poster
Technical Officer: Luca Ferracina ([email protected]) 49 Prototyping a system for parachute dynamic extraction test
Prog. TDE Budget: 800 KE Duration: 24 months Aim: D Ref.: 4971 (STAT) TRL: 2 to 3
Objective ExoMars Dynamic extraction tests To prototype a system for parachutes dynamic extraction test
Description In this activity a system for parachutes dynamic extraction tests will be prototyped to allow preliminary design verification and performance of parachute bag design and parachute integrity during extraction. The objective of this activity will be also to test the system in representative conditions. This is very relevant for ESA missions that require a pilot chute deployment system (opposite to mortar deployment) like the one used in ExoMars and preliminary based line for M* (Ice Giants) mission or similar science or exploration missions e.g. on Titan, Venus, etc. The activity is composed of the following major tasks: 1. review of the present methodologies for testing parachute deployment using a pilot parachute. The overview should cover (but not limit to) systems using a (balloon / aircraft) drop test, pulley system, a catapult, a tractor rocket system, bungee and a “car- pulling” system. New systems shall be proposed if necessary. 2. trade-off and selection of the most suitable and economically affordable system for a typical application (to be coordinated with the Agency) and associated conditions 3. design the selected system and define a development plan including the tests to be executed 4. (design if required and) procure a parachute system to be tested 5. Perform all activities described within the development plan, hardware manufacturing as well as integration procurement/manufacture of instrumentation system, integration and system calibration, perform any functional test (if required) 6. execute the preliminary testing with the new developed system using the procured parachute Low Density Supersonic 7. leveraging on the results obtained during the test campaign, a complete development roadmap will be defined to further Decelerator Parachute improve the proposed system. Decelerator- AIAA 2013-1329 Background and previous activities System The concept and approach to technologies is applicable to any science/exploration missions requiring a parachute deployment using a pilot chute for the extraction (e.g on Mars Titan, Venus, etc.) Activity have a strong link with the Mars Future Exploration Mission Scenarios.
Technical Officer: Luca Ferracina ([email protected]) 50 Preliminary Design and Performance verification for of critical elements for high precision landing on Mars Prog. TDE Budget: 800 KE Duration: 24 month Aim: D Ref.: New TRL: 2 to 4 Objective(s) Preliminary design, prototype and testing of a Reaction Control System (RCS) for a Mars precision landing.
Description Entry thrusters can have 2 purposes for an EDL system : attitude control or entry guidance. In both cases, aerodynamics and aerothermal interactions of RCS thrusters with the flow at hypersonic velocities (or supersonic) and the probe itself need to be analyzed well ahead of the mission; plume interaction can for instance generate a local heat increase on the probe surface, and jet interaction can cause un-desired forces / torques, which could jeopardize the controllability of the probe.
A similar activity has been conducted but not completed for programmatic reasons in 2015-2017. The present activity shall revisit the already obtained results and concentrate on the following objectives: 1. preliminary design of a RCS capable to control and guide (for a precision landing) an entry probe on Mars; 2. derive the requirements for a test campaign (in an aerodynamics rarefied facility with high velocity and low pressure) to verify the performance of the designed system; 3. prepare (and eventually modify/upgrade) the selected facility and measurement techniques where to perform the test campaign. 4. Procure / manufacture the required hardware 5. Test the Reaction Control System
Background and previous activities A similar activity has been conducted but not completed for programmatic reasons in 2015-2017
Preliminary Design and Performance Verification of a Guided Technical Officer: Luca Ferracina ([email protected]) Entry Thruster System for Precision Landing on Mars - EUCASS2017-261 51 EDL for Mars penetrators
Prog. GSTP Budget: 500 KE Duration: 24 Aim: D Ref.: New TRL: 3 - 5 Objective(s) Aerothermodynamics assessment of a possible Mars penetrator
Description Within the present activity an aerothermodynamics assessment of a spiked/bullet shape entry configuration will be performed by analyzing the gas chemistry during the entry, the resulting heating and the required thermal protection systems, the required verticalization of the trajectory prior the impact and the associated drag and stability performance of the braking system, the effect of the wind and the controllability of the impact conditions. The activity shall complete the analysis with dedicated tests to confirm the suitability of the propose solution.
Background and previous activities A preliminary study (Aerothermodynamics assessment of an Entry Descent and Landing system for a penetrator mission) has been conducted in 2018 – 2018. There have been several attempts to deploy a penetrator system into the surface of a planetary body to investigate its surface and interior: Deep Space 2 microprobes, Lunar-A and Mars 96. Although none of these missions have been successfully completed, a penetrator remains an attractive option for future exploration missions. As also recently pointed out in recent penetrator studies, for missions to bodies with atmospheres such as Mars, a relatively high drag area is needed to get the required verticalization of the trajectory prior the impact . The simplest proven way to achieve this would be to enclose the penetrator in a capsule. Although this solution takes advantage of existing knowledge and capability in Europe on conventional capsule performance, it requires a large volume and consequently a rather high mass. A spiked/bullet entry shape approach could potentially reduce the required space and mass needed.
Courtesy of CFS Technical Officer: Luca Ferracina ([email protected]) 52 Day 2
53 Computational Fluid Dynamics Acceleration Through Hardware
Prog. GSTP Budget: 500 kEUR Duration: 12 months Aim: A Ref.: A09 TRL: 4 -> 5
Objective(s)
To investigate, develop and implement techniques to accelerate computational fluid dynamics by means of GPU and/or reconfigurable hardware. To apply such hardware acceleration techniques to conventional available CFD solvers To decrease the computational time by at least one order of magnitude Description High performance computing is advancing at the same pace as the numerical algorithms increase complexity. In order to shorten the design iterations during concurrent design and to speed up troubleshooting by analysis, a fundamental change of the computational paradigm is required. Development of novel optimum algorithms has always been one way towards this end, the other being the advances in hardware, like multiple core / multiple CPUs architectures which speed up the simulation by means of job parallelization. General purpose GPU have enabled a significant speed up but oftentimes this comes at the cost of tailoring of the algorithms, thus making this technology only available to some specific methods, e.g. lattice Boltzmann. On the other hand, the use of reconfigurable hardware like FPGA has been suggested or even applied to numerical solvers of fluid dynamics. In this case, the potential speed up may largely trade-off for the added complexity of a simultaneous coding by hardware and software, and may eventually enable real time CFD. The activity shall comprise: - The implementation of hardware acceleration techniques to conventional available CFD solvers. - Time accurate and steady state CFD solvers shall be targeted.
Background and previous activities
Previous collaborative works between INTA and DLR (not ESA funded) have looked into the application of FPGA programming to the CFD code TAU.
Vulcain 2.1 for Ariane 6 during acceptance testing. © ArianeGroup
Technical Officer: Victor Fernandez Villace ([email protected]) 54 Measurement Techniques for Storable Propellant Tanks
Prog. TDE Budget: 400 kEUR Duration: 12 months Aim: C Ref.: C02 TRL: 2 -> 3
Objective(s)
To design a propellant gauging method based on the response to mechanical excitations To demonstrate by analysis the performance To demonstrate by experimentation the performance To assess the feasibility against conventional gauging methods
Description
Conventional propellant gauging methods e.g. PVT, thermal knocking or bookkeeping suffer from considerable uncertainties due to heat loss, pressuring gas absorption, error propagation. The next generation of exploration missions beyond LEO require a reliable propellant gauging capability.
The proposed activity shall comprise the following tasks: - Analytical demonstration of the concept - Breadboard demonstration of the critical functions and performance - Assessment of the performance of the envisaged flight system in comparison to conventional techniques
Background and previous activities
When subject to mechanical excitations, the frequency response of the tank structure is dependent on the amount of liquid in contact with the tank. The spectral characterization of the vibrations sensed on the tank wall serves to identify modal changes based on the fill level. This methodology thus results in an absolute determination of the propellant mass which eliminates the main sources of error associated to the classical indirect methods. Propellant tank of the European Service Module-2 © Airbus
Technical Officer: Victor Fernandez Villace ([email protected]) 55 Validation of Complex Propellant Network Models
Prog. GSTP Budget: 600 kEUR Duration: 12 months Aim: D Ref.: - TRL: 4 -> 5
Objective(s)
To demonstrate an accuracy of the analytical model above 10%
Description
There exist numerical tools to assess the performance of complex propellant networks when subjected to the highly transient regimes encountered during start up and operation of spacecraft propulsion systems, e.g. priming and cross talk. The analysis tools are of high fidelity but need to be validated against representative test cases. The proposed activity will validate the analyses by comparing the predicted fluid pressure and temperature against the measured data for saturation conditions of the liquid which are relevant to the propulsion system. Highly accurate pressure measurements will be needed to validate the high frequency signature (amplitude and damping) of the predicted pressure signal.
The activity shall comprise the following tasks: - Development of the hydraulic test bench: test facility with propulsion-system representative components (pipes, valves, tanks, manifolds...) and instrumentation (time resolved pressure, temperature and mass flow) - Test campaign: elaboration of the test matrix and conduction of the tests and post-processing of the results. - Numerical campaign: numerical rebuilding of the test cases and validation against the test data, refinement of the numerical tool
Background and previous activities
The activity builds upon the existing capabilities of the European Space Propulsion System Simulation (ESPSS), which comprises a set of high fidelity analytical models of propulsion components that enable the simulation of propellant networks. Rear view of the European Service Module-1 © ESA–A. Conigli
Technical Officer: Victor Fernandez Villace ([email protected]) 56 Pneumatic-Thermodynamic Models
Prog. GSTP Budget: 500 kEUR Duration: 12 months Aim: A Ref.: A21 TRL: 4 -> 5
Objective(s)
- To develop high-fidelity dynamic models of pressurization systems - To validate the models with experimental data
Description
Propulsion pressurization performance models shall be improved by implementation of link with thermal models, the expected activities should include representative testing to check/update the theoretical models and related parameters.
The activity shall encompass: - The development of analytical models of pressurization systems - Verification against experimental data
Background and previous activities
There is a need for accurate multidisciplinary (thermal, mechanical, fluid dynamic, electrical) models of propulsion pressurisation system and valve components. These models allow a more efficient preliminary design phase and accurate assessment and troubleshooting of the performance at component and system/assembly levels. Reaction control system of the Exomars Orbiter © ArianeGroup
Technical Officer: Victor Fernandez Villace ([email protected]) 57 Demonstrator of a Detonation Propulsion System
Prog. GSTP Budget: 600 kEUR Duration: 24 months Aim: D Ref.: - TRL: 3 -> 4
Objective(s)
To demonstrate the thrust and Isp performances of a detonation propulsion system
Description
The activity shall design and test a complete pulse detonation system composed of the propellant storage assembly, the propellant feed system, the detonation driven combustion chamber, the expansion nozzle.
This activity encompasses the following tasks: - Development and/or validation of system level analytical models - Design and manufacture of the detonation propulsion system - Testing of the detonation propulsion system
Background and previous activities
Pulsed detonation engines (PDE) exhibit a superior performance in terms of specific impulse than conventional liquid, hybrid or solid rocket engines. In addition, the pulsed detonation combustion offers the advantage of lowering the required injection pressure. This allows lowering the requirements on the pressure regulation system needed for the current rocket engines based on a constant pressure combustion. Because of this, the PDE can potentially simplify and reduce the weight of the current propulsion systems. Although the research on PDE started more than half century ago, the current modelling capabilities and technology justify revisiting this concept. Simulation of the detonation front in a detonation tube Follow up of “Pulsed Detonation Thruster” © von Karman Institute for Fluid Dynamics
Technical Officer: Victor Fernandez Villace ([email protected]) 58 Assessment and Improvement of Turbo-pump Chill-down
Prog. GSTP Budget: 300 kEUR Duration: 24 months Aim: D Ref.: - TRL: 4 -> 5
Objective(s)
To develop higher fidelity engineering models to assess the chill-down of cryogenic propellant pumps. To validate the engineering models with chill-down data obtained from representative tests. To minimize the wasted propellant optimizing the pump design and chill-down process.
Description
The chill down of cryogenic propellant pumps typically wastes several tens of kilograms. A reduction of one kilogram of liquid oxygen allows saving tens of thousands of euros due to the extra payload capacity gained. The design of the chill down sequence is currently tackled by means of low fidelity engineering models and is heavily supported by expensive experimentation in cryogenic facilities. Aided by higher fidelity models, the optimization of the pump design and chill down process will allow to reduce the amount of wasted propellant.
The activity shall encompass the following tasks:
- Development of fluid to solid heat transfer models - Experimental validation of the heat transfer models - Optimization of the pump chilling process
Background and previous activities
The start-up sequence of a liquid rocket is an extremely critical phase which exerts highly transient thermo-mechanical loads on the engine components. In particular, the turbomachinery of cryogenic liquid stages are subject to extreme thermal gradients and dynamic loads which, if not properly managed, can lead to a catastrophic failure of the components. In order to prevent this, the engine start-up sequence begins by chilling down the turbomachinery to cryogenic temperatures. This allows to limit the thermal gradients encountered during engine start, prevents the appearance of cavitation and ensures Artistic rendering of the LOX-LCH4 Prometheus engine that the operational conditions, e.g. clearances and pre-loads, of the components are reached prior to rotor speed up. © ArianeGroup
Technical Officer: Victor Fernandez Villace ([email protected]) 59 Aero-thermodynamics characterization of re-entry bodies in hyper-velocity flows
Prog. GSTP Budget: 500 kEUR Duration: 12 months Aim: D Ref.: - TRL: 4 -> 5
Objective(s)
To design and implement a test section for the ESTHER shock tube
Description
The European Shock Tube for High Enthalpy Research (ESTHER) lacks of a test section for the aero-thermal characterisation of entry shapes. The experimental measurement of the aero-thermodynamic characteristics of entry bodies allows the verification of the design methodologies on three-dimensional geometries.
The activity shall comprise the following tasks:
- Design and provision of a test section to enable testing of aero-shapes - Design and provision of instrumentation to characterize the free stream flow - Experimental campaign and commissioning of the test section
Background and previous activities
The ESTHER shock-tube implements a set of novel technologies, which enable a key set of specifications and requirements which are critical for a new generation facility in support to the European ambitions for planetary exploration in the 21st Century, namely: reliability, repeatability, safety, high performance, cleanliness, low-cost, high turnaround. ESTHER shock tube in the Institute for Plasmas and Nuclear Fusion in Lisbon © ESA & IST
Technical Officer: Victor Fernandez Villace ([email protected]) 60 Development of a Step and Gap Methodology for AIV/AIT from an Aerodynamic Perspective Prog. GSTP Budget: 300k€ Duration: 24 Aim: A Ref.: 14 TRL: 3 to Target 4 Objective(s):
Establishing engineering design rules for acceptable tolerances during the AIV/AIT phase of flight vehicles to avoid overheating and pressure losses.
Description:
The proposed activity is to experimentally study irregularities in geometries such as steps, gaps, seals etc... The outcome should allow establishing a methodology and related correlations for irregularities which are critical for real applications. These will serve as a guideline for the definition of more accurate safety margins and related tolerances for AIV/AIV operations. These correlations shall give the designer the tools to still assure an aerodynamically smooth surface despite the presence of protuberances or imperfections. The complex aerodynamic and heat transfer effects in the region of steps and gaps will be characterised by applying an array of measurement image techniques during the same test. This will lead to the development of correlations enabling the determination of critical dimensions to assure an aero(thermo)dynamically smooth surface without local overheats, sneak flows or triggering any other adverse flow phenomena.
Background and previous activities:
- Completed de-risk activity to accommodate various protuberance geometries in one single WT model
Technical Officer: Johan Steelant ([email protected]) 61 Morphing of Propulsive Aero-Thermally Loaded Structures
Prog. TRP Budget: 600k€ Duration: 24 Aim: A Ref.: NEW TRL: 3 to 6
Objective(s):
• Exploring the potential of adapting the geometry of elements in a propulsive flowpath exposed to high aero- thermal loads. • Developing a demonstrator to adapt a aero-thermal loaded geometry during operation.
Description:
• During the high-speed part of an ascent or descent trajectory, the geometry of the inner flowpath needs to be adapted to optimize the operation of the propulsion unit. This could e.g. be a variable intake or nozzle,... This morphing capability would have the additional advantage that gaps or cavities e.g. around intake panels, nozzles... are not any longer present and hence local hot spots and sneak flows can be avoided.
• The activity will develop a demonstrator which structure consists of a high temperature resistant (non)- metallic materials and can adapt its geometry during operation in a high temperature flow.
• This demonstrator could represent e.g. an adaptable nozzle or intake throats, compression or expansion contours, etc.
Background and previous activities
Technical Officer: Johan Steelant ([email protected]) 62 Disruptive Conceptual Sizing and Cost Optimization Methodology of Reusable Flight Vehicles Prog.: GSTP Budget: 800k€ Duration: 18 Aim: A Ref.: 16 TRL: 3 to Target 6
Objective(s):
• The first objective of the activity is to assess mass, volume and power budgets of an expandable and reusable vehicle and its on-board system as from the conceptual design phase. • The second objective is to assess the related cost for the overall concept in terms of recurring and non-recurring costs.
Description: This activity targets both space exploration and space transportation systems in the areas of re-usability: • Space exploration: re-usability concepts are required in the HERACLES exploration activities where re-usability of a landing stage on the South Pole of the Moon is a key feature. • Space transportation systems re-usability of launch stages: key feature for launch price reduction.
The research goal will be pursued through the implementation and exploitation of ad-hoc surrogate models per each main equipment. Surrogate models are mathematical relationships, which express mass and volume as functions of physical and performance characteristics. Surrogate models will be based on the combination of results coming both from functional simulations and current statistical data.
The economic sustainability shall be considered in this overall process, not only from the point of view of development and production costs but also from the operation and exploitation perspective. Appropriate cost models shall be implemented able to cover development and operational costs, (non)-recurrent costs considering both expandable and reusable applications.
Background and previous activities:
• Some feasibility studies performed.
Technical Officer: Johan Steelant ([email protected]) 63 Transient Multi-Phase Engineering Assessment Tools for Cryogenic and Green Propellants Prog.: TRP Budget: 650k€ Duration: 18 Aim: A Ref.: NEW TRL: 3 to Target 5
Objective(s):
Development of a generic engineering tool to assess the transient behavior of various multi-phase phenomena occurring during the operational changes of a propulsive system.
Description:
The overall design of propulsion systems is mainly dictated by the transient occurring during the start-up and shut- down of the rocket engines, whether it be for spacecraft or launchers (expandable or reusable). The transient is mostly characterized by multi-phase phenomena typically occurring during these operational phases of the engine: - Exposure to vacuum conditions resulting into sudden phases changes: flashing, cavitation… - Exposure to a hot or cold engine resulting into explosive boiling, vapour locks, … - Exposure to micro-gravity where multi-phase phenomena are distinctly different to what occurs in ground facilities - Exposure to sudden operational changes of the propulsion system results into highly unsteady behaviours such as waterhammer, instabilities, acoustic reflections
Describing all these phenomena in a generic engineering assessment tool requires not only the modelling of the different phases but also their strong interaction. Recent development indicate now the available maturity to incorporate this into design and assessment tools for propulsion systems for various types of cryogenic and green propellants. This modelling capability will be incorporated in the generic open-source engineering tools for propulsion system supported by ESA.
Background and previous activities:
Technical Officer: Johan Steelant ([email protected]) 64 AIV/AIT Impact on Aero-Thermo-Mechanic Load Uncertainty Quantification Prog.: TRP Budget: 500k€ Duration: 18 Aim: A Ref.: NEW TRL: 3 to Target 6
Objective(s): - establish a methodology to quantify the uncertainties and define allowable tolerances during detailed design, manufacturing and AIV/AIT operations - establish a set of correlations based on dedicated experimental campaigns enabling the quantification of allowable tolerances and uncertainties
Description:
The proposed activity is to experimentally study irregularities in geometries such as protuberances or at interfaces (e.g. steps, gaps, seals etc...). The outcome shall allow establishing a methodology and related correlations to assess the occurrence of criticalities and their locations for real applications. These shall result in guidelines for the definition of more accurate safety margins and needed tolerances during detailed design and AIV/AIV operations. These correlations shall give the designer the tools to still assure physically a smooth surface despite the presence of protuberances or imperfections with little or no impact on the aero-thermo-structural performance of the vehicle.
The complex aerodynamic and heat transfer effects in the region of interfaces/protuberances will be characterised by applying an array of measurement techniques during the experimental campaigns. This shall lead to the development of correlations enabling the determination of critical dimensions to assure from a macroscopic point of view a smooth surface without local overheats, sneak flows or triggering any other adverse flow phenomena.
Background and previous activities:
• .
Technical Officer: Johan Steelant ([email protected]) 65 Multi-disciplinary Design Methodology for Quiet High- Speed Wind tunnels Prog.: TRP Budget: 600k€ Duration: 18 Aim: B Ref.: NEW TRL: 1 to Target 2
Objective(s):
- Setting up a design methodology to properly design a quiet wind tunnel for high-speed vehicle design on the basis of various expert tools - The generic methodology shall be also applicable to the design of future reusable launch and re-entry vehicles.
Description:
The design of a quiet nozzle demands the seamless combination of various expert tools enabling: - Correct contouring of a nozzle wall to assure a uniform exit flow - Correct prediction of the boundary layer and their correction on the main flow with particular attention to flow instabilities induced by acoustic, entropy and velocity perturbation in the mean stream - Enabling to delay the boundary layer growth by means of bleeding or suction technologies without introduction of parasitic perturbations - Mastering the thermal aspect on the instabilities and the overall dilations of the nozzle geometry - Mastering the roughness requirements and their effect on the perturbation
Background and previous activities:
Technical Officer: Johan Steelant ([email protected]) 66 Magnetic Suspension Balance Systems (MSBS)
Prog.: TRP Budget: 500k€ Duration: 24 Aim: B Ref.: NEW TRL: 1 to Target 3
Objective(s):
• making sting-free measurements of time-dependent forces acting on aerodynamic models. • to identify unsteady aerodynamic phenomena that are of paramount importance for control of highly manoeuvrable vehicles operating at high (post-stall) angles of attack
Description: an MSBS offers the best way of estimating forces on engineering bodies as in ‘flight’, such as launchers, landers, drones and re-entry space capsules. Often such bodies operate in a mixed ballistic / aerodynamically assisted regime. The understanding of the fluid forces is important when the recovery of a launcher stage or space capsule depends on an accurate measurement of the operating forces, lift and drag. Measurement of dynamic properties (mass, and inertia of the vehicle) and aerodynamic properties (of the vehicle and its control surfaces) are key to precise control. MSBS offers the opportunity to identify unsteady aerodynamic phenomena that are of paramount importance for control of highly manoeuvrable vehicles operating at high (post-stall) angles of attack or indeed novel vehicle configurations and flapping flight. Hardware-in-the-loop simulations would enable free-flight testing of such vehicles in the wind tunnel.
The activity would entail a feasibility study where the focus would be related to the active control definition and implementation of a generic MSBS, while respecting the requirements from the different envisaged mission
Background and previous activities:
• N/A: only available at JAXA and Tohoku university..
Technical Officer: Johan Steelant ([email protected]) 67 Combined Aero-Propulsive Testing Capabilities with Active Rocket Engines Prog.: GSTP Budget: 500k€ Duration: 18 Aim: B Ref.: NEW TRL: 3 to Target 5
Objective(s):
- Development of an automated control for the aero-propulsive testing enabling - Assessing feasibility of retro-propulsive mode for RLV and landers including surface control devices
Description:
The feasibility study aimed at demonstrating that a propulsive unit with a separated balance can be incorporated within the geometrical constraints of a launcher geometry. The aim was mainly to assess the base pressure during ascent with engine(s) on for a single operational point.
First aim for present activity is to have an automated control of the propulsive units independently from the aerodynamic operation of the tunnel. This requires a control logic as well a dedicated feeding system for the propulsion unit.
So far the feasibility was limited to a classical operation of an expandable launcher during ascent. The technical challenge increases when considering retro-propulsive testing at various Mach numbers from subsonic up to Mach 4. The model shall then also entail aerodynamic flight controls to characterize fully the aero-propulsive behavior of a RLV, needed for future GNC developments.
Background and previous activities:
• Building upon the experience gained in the feasibility study by INCAS
Technical Officer: Johan Steelant ([email protected]) 68 Bio-Inspired Micro-Textures on Improving Flight Control Devices Prog.: TRP Budget: 500k€ Duration: 18 Aim: B Ref.: NEW TRL: 3 to Target 6
Objective(s):
Assessing the implementation of bio-inspired micro-texture on improving the efficiencies of flight control devices
Description:
Various types of bio-inspired micro-textures have also their benefit on high-speed flows in attenuating unsteadiness and instabilities encountered in the flow. This allows the postpone particular negative effects such as flow separation or delaying the growth of unwanted instabilities. This has its benefit in an overall lower drag, improved flight control and widening the operational flight domain of launchers and re-entry vehicles. The activity will pursue various ways to implement bio-inspired textures enabling the postponement or elimination of particular flow phenomena such as encountered in adverse pressure gradients or instabilities growth in boundary layers. These are typical on ramps and flight control devices resulting in increased efficiencies and lowered thermo- mechanical loads. The most promising results will be experimentally tested and supported by numerical investigations to assess the fundamentals of the increased performance.
Background and previous activities:
• .
Technical Officer: Johan Steelant ([email protected]) 69 Development of a Hypersonic Flight Research Vehicle Demonstrator Prog.: GSTP Budget: 1000k€ Duration: 18 Aim: D Ref.: 4 TRL: 3 to Target 6
Objective(s):
Development of a generic service module to lower the cost and to ease the execution and preparation towards various types of flight experiments.
Description: Development of various technologies for high-speed vehicles are hampered by the operational limits of on-ground facilities. This covers various aspects such as limited size or operational time in high-speed wind tunnels, not representative flight conditions such as free stream turbulence, heated up walls, flight stability, GNC and control. Certain techniques, methodologies and/or hardware need to be tested during free-flight prior to be installed on launchers, (re)entry vehicles. Flight experiments for these kind of technology demonstration have already been shown to be very successful by using a flying platform on the basis of a modified sounding rocket enabling suppressed trajectories. So far, modifications and qualification of these sounding rockets were always customised for each flight experiment. It has been shown though that many elements can be made generic and standardised to minimise time and cost efforts through the development of a generic service module and related peripherals enabling various mission requirements. The present study aims to gather the experience and specific needs from various experimental fights to identify the common grounds and define the general requirements. This shall lead to a detailed design of the various components which are classically adapted, devised for each mission.
Background and previous activities:
• .
Technical Officer: Johan Steelant ([email protected]) 70 Prototype of Re-usable Concepts for Space Transportation Flight Vehicle Engineering Prog.: GSTP Budget: 6000k€ Duration: 18 Aim: D Ref.: 8 TRL: 3 to Target 6
Objective(s): Phase B of a Hypersonic Flight Research Vehicle Demonstrator that shall allow the development of key functionalities of hypersonic flight technology capabilities in the following areas: • Advanced propulsion functions for hypersonic flight • Aero thermodynamics shapes for variable flight regimes • Versatile flow control • State of the art sensors and health management for hypersonic flight • Thermal balance and heat management • Durable materials and light structures for hypersonic flight
Description:
The activity would comprise 2 project reviews associated with the Phase B.
1. The system requirements review (SRR) held during the course of Phase B with the aim to release updated technical requirements specifications, to assess the preliminary design definition and to assess the preliminary verification program. 2. The preliminary design review (PDR) held at the end of Phase B. The outcome of this review will be used to judge the readiness of the project to move into Phase C.
Background and previous activities:
• .
Technical Officer: Johan Steelant ([email protected]) 71 Development of an Experiment for Transient Phase Change Aspects in Microgravity for Storable or Cryogenic media Prog.: GSTP Budget: 600k€ Duration: 24 Aim: D Ref.: 13 TRL: 4 to Target 7
Objective(s):
Development of a dedicated experimental facility to address multi-phase flow phenomena and phase changes under microgravity conditions
Description:
Various missions are faced with phase change on board of vehicles with different applications: - cooling of scientific (sensors, optics…) by means of liquid helium - cooling of on-board systems of spacecraft (transponders, …) by means of phase change fluid loops - recommissioning of upper stages after coasting phases (chill-down processes....).
All of them have in common that the phase-change and heat transfer is behaving completely different on Earth and microgravity. Most commonly the pressure drop for multi-phase flow under microgravity is substantially higher. Phase change is also more difficult to realize in microgravity but can be enhanced in certain circumstances or for certain applications. In order to assess these phenomena, dedicated microgravity experiments are needed for various fluids to assess their effect during multi-phase and phase change operations. A dedicated experimental module will be devised to be flow on existing sounding rocket opportunities (e.g. Rexus). This will include also the related modelling of Multi- Phase flow with Phase change under slow and fast transient conditions in classical 1D engineering fluid models.
Background and previous activities:
• .
Technical Officer: Johan Steelant ([email protected]) 72 Atmospheric Mapping by Automated Probe Drones on Skipping Trajectories Prog.: TRP Budget: 500k€ Duration: 18 Aim: D Ref.: NEW TRL: 2 to Target 4
Objective(s):
Feasibility study to map the atmosphere by automated probe drones released from a spacecraft in orbit Demonstration of a prototype/breadboard with integrated systems on a ballistic ground facility
Description: Free flight testing of drones at subsonic speed in combination with active control and auto-piloted navigation is presently commonplace in the aeronautical world. Particularly the availability of miniaturized components such as actuators, IMU, control surfaces lowered drastically the realizibility threshold. However, free-fly testing a scaled concept at supersonic and hypersonic speeds was not yet demonstrated so far. This high- speed probes would be released from an orbiting spacecraft at high- speed and enter any type of planetary atmosphere through a controlled skipping trajectory. During these passages through the atmosphere, the drone would contain probes to map the atmosphere locally with respect to the scientific needs: composition, temperature, pressure… The activity would require the manufacturing of a prototype high-speed lifting vehicle including the integration of the electronics (IMU, GNC, onboard power), mechanics (e.g. actuators, control surfaces…), telemetry and data acquisition while being exposed to high temperatures and acceleration forces.
Background and previous activities:
Technical Officer: Johan Steelant ([email protected]) 73 Simulation of Spacecraft Generated Lunar Exosphere and Global/Local Ground Contamination Prog.: TDE Budget: 350 k€ Duration: 12 months Aim: A Ref.: NEW TRL: 2 - 4
Objective This activity shall help to assess and plan future scientific missions on the Moon. In this context a numerical tool shall be developed, enabling to estimate both the local and global distribution and evolution of volatiles, originating from a vehicle’s propulsion system during landing.
© ESA Description Any powered landing on the lunar surface will release a large amount of volatiles. Due the Moon’s lack of an atmosphere and it’s rather low gravity, the propagation of exhaust gas traces is not limited to the surroundings of the vehicle or the local landing site. Studies and simulations have shown that exhaust gases can disperse on a global scale and can temporarily or © ESA permanently contaminate large parts of the surface and thus might interfere with scientific experiments conducted by humans, lander or rovers.
Within this activity a numerical tool shall be derived, allowing the assessment and simulation of several critical aspects:
• Simulate the evolution and propagation of critical species (e.g. H2O), which are released during the descent of the lander (e.g. using an adapted DSMC code). The landing site and time, the descent trajectory and attitude, as well as engine characteristics shall be modifiable. • The interaction (e.g. adsorbed, permanently trapped, photodestroyed) of the released species with the lunar surface shall be simulated for periods up to a few days after landing
Background and previous activities The search for and detection of water is one of the mission goals envisaged for EL3 (European Large Logistics Lander). In order to plan the missions (e.g. landing site) and surface operations (e.g. pathway of rover) a good understanding of the exhaust gas behavior and cloud evolution is crucial.
The Evolution of a Spacecraft-Generated Lunar Exosphere, Prem et al., 2009
Technical Officer: Stephan Schuster ([email protected]) 74 Methods for Characterisation of Plume Radiation
Prog.: GSTP Budget: 600 k€ Duration: 18 months Aim: C Ref.: C07 TRL: (current to Target)
Objectives The focus of the work shall be the development and application of spectrometers and radiometers for the characterization of rocket plume flows in a ground facility. In addition to the systematic experiments with advanced measurement techniques, a numerical study is needed to compare the results and to identify necessary improvements in modelling and numerical simulation.
Description The base of launchers (with multiple nozzles) is exposed to a highly non-stationary flow field, which partially contains separated and vortex embedded high temperature flow regions. Depending on the propulsion type and the gas chemistry, the related gas radiation may vary strongly. Although the convective base heating at high flight altitudes is weak, the radiative heating of the gas is still significant. This has a direct impact on the design margins of the base structure and its TPS. The prediction of the radiative heat transfers is still the weak point of computational design tools. Therefore a reliable experimental database needs to be generated to close this gap. In addition, a numerical study shall be performed to and compared with the experimental results. This should help to improve the simulation modelling and reduce uncertainties in the prediction. Since the number of flight experiments is limited, only systematic ground tests using advanced diagnostics would provide such a data set. In the long term these ground qualified measurement techniques should be used in real flights. Radiometers, which will be developed within this study, should be qualified for a potential flight on of the European launchers by means of passing mechanical tests, thermal tests, shock tests and EMC tests of an selected mission.
Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (C07)
© NASA – Thermal and Fluids Analysis Workshop 2017
Technical Officer: Stephan Schuster ([email protected]) 75 Flush Air Data Sensing for Entry Missions
Prog.: GSTP Budget: 800 k€ Duration: 24 months Aim: C Ref.: C08 TRL: (current to Target)
Objectives Development and testing of a Flush Air Data Sensing (FADS) system for entry missions. The sensor shall cover low speed entry (LEO) and high speed entry (Mars, return to Earth).
Description Flush Air Data Sensing (FADS) makes use of surface pressure measurements from the nose cap (or the leading edge) of a vehicle and uses this data to derive real-time atmospheric- and flight parameters like angle of attack, side slip angle, Mach number and altitude. © NASA X-43 Project
This information can be fed to the on-board guidance and control system of the vehicle for in-flight corrections of the re-entry © DLR Cologne attitude, as well as for flight data recording and post flight analyses.
Within this activity a Flush Air Data Sensing system shall be developed (including sensors, software algorithms) and shall be tested and validated for the application in low and high speed re-entry scenarios.
Background and previous activities Activity is captured in the Technology Harmonisation Dossier 2020 on ‘Fluid Mechanics and Aerothermodynamic Tools’ (C08)
In-Flight Demonstration of a Real-Time Flush Airdata Sensing (RT-FADS) System, Whitmore et al., 1995 Technical Officer: Stephan Schuster ([email protected]) 76 SmallSat Exploration Missions Utilizing Commercial European Microlauncher Capabilities Prog.: TDE Budget: 400 k€ Duration: 12 months Aim: D Ref.: NEW TRL: 2 - 4
Objectives The objective of the activity is the assessment of the beyond low Earth orbit (LEO) performance of European microlauncher (concepts), the identification of possible mission enabling or performance increasing launcher upgrades (e.g. kick-stages) and the proposal of technically feasible and scientifically viable mission concepts.
Description A significant number of commercial microlaunchers are currently under development within Europe (HyImpulse, IsarAerospace, PLD Space, Orbex, RFA, …), with the goal of being operational and offering launch services within the next few years. They are primarily designed for the delivery of small payloads (<1.5 tons) into low Earth orbits. In principle however, they can also be utilized for exploration mission to destinations beyond LEO (e.g. Moon).
Within this study, European microlauncher systems shall be analyzed in terms of beyond-LEO performance. Critical points shall be identified and necessary or possible launcher upgrades shall be highlighted.
Various scenarios, including different target bodies (e.g. Moon), transfer strategies and arrival conditions (e.g. fly-by, hard landers, in- orbit insertions) shall be evaluated and technically feasible and scientifically viable mission architecture concepts shall be proposed.
A preliminary design of realizable payloads (e.g. CubeSats, scientific instruments, cameras) shall be drafted and the overall mission cost (incl. dedicated launch procurement) shall be evaluated.
Background and previous activities Large amount of activities in TEC-MPA related to conceptualization and assessment of microlaunchers. Support to the Boost! Programme (ESA’s Commercial Space Transportation Services and Support Programme).
© Rocket Lab
Technical Officer: Stephan Schuster ([email protected]) 77 Efficient Meshing and Post Processing for Aerothermodynamics Computations Prog: GSTP Budget: 700K 18 months Aim: A Ref.: 10 TRL: 4 to 5 Objective(s): Automation of 3D meshing and post-processing of a complex geometries optimizing the pre- and post processing activities for ATD CFD simulations.
Description:
A significant amount of the manpower in simulations is spent cleaning the CAD files, making it watertight, and removing or simplifying local geometries in generating meshes. This activity will aim to reduce the overall time generating a mesh, and easy the postprocessing of the large datasets generated on distributed HPC clusters.
The activity shall include the following: 1. A unified user interface plugged-in CFD code and different types of mesh generators and CFD codes. 2. Basic functionalities, such as import of at least STEP surfaces, adding or deleting local geometries, closing holes and able to clean-up the surface geometry shall be implemented. 3. The solution shall at least cover the most commonly used types of meshes: structured multi-block, overset, unstructured tetrahedral, unstructured tetrahedral + prisms for the boundary layer, unstructured tetrahedral + hexahedral for the boundary layer, unstructured polyhedral, Octree 4. An integrated post-processing tools for aerothermal applications allowing: • Integration of Volume Streamlines, Surface streamlines, Slices, iso-surfaces, Zone extraction, numerical schlieren, shadow graph • Interpolation of solution on different mesh types, • Automatic report generation (standardized output) Background and previous activities:
In the today’s ATD community a large number of computational fluid dynamics codes are used. Three types CFD codes exists: commercial, institutional, and open source. While the commercial codes include a user friendly interface for mesh generation and post-processing, the latter two codes are often interfaced with open source or commercial mesh generation and post-processing software. Most of the manpower in performing simulations lies in the pre processing and the handling of the large dataset for postprocessing. A unified user friendly GUI for the different ATD codes will increase the efficiency performing the simulation and allows running several solvers on exactly the same setup and exchange and compare the data sets.
Technical Officer: Louis Walpot ([email protected]) 78 Development of advanced state-to state thermo-physical models for high energy outer gas planet entries Prog: TDE Budget: 300K Duration: 18 months Aim: B Ref.: 4 TRL: 4 to 5
Objective(s): The goal of this research activity is the development of a high fidelity, high-speed state-to-state, 2D/3D numerical tool for the reliable determination of the heat fluxes during entry into gas planetary atmospheres. The numerical tools shall implement the most recent and accurate dynamical data for elementary processes and exploiting the efficiency of modern computing systems. Description: The purpose of the project shall be to build a Navier-Stokes code with state-to-state kinetics for outer gas planets atmospheric entry for massively parallel calculation, and to validate the data with experimental results when available, so as to provide an effective virtual wind tunnel for high enthalpy and entry flows. The fluid dynamic code shall be built in the two- and three-dimensional versions and equipped with state-to-state non-equilibrium models. Also, the effective time-dependent entry problem could be faced with the vehicle being moved while interacting with the flow without the need for expensive mesh re-generation. An existing code for radiation modelling shall be used to estimate the radiative heat flux, updating, when necessary, the database of emission properties and collisional data. The following detailed tasks are foreseen: T1 - Implementation of the code the ionization kinetics for H2/He/CH4 plasmas. T2 - Implementation of discretization routines, based on multi-block body-fitted structured grid and immersed-boundary technique, in the CFD code T3 - Refinement of state-to-state chemical model for H2/He/CH4 plasmas with impurities T4 - Calculation of thermodynamic and transport properties of outer gas planet atmospheres contaminated by species ablated from TPS tiles T5 - Construction of macroscopic model for outer gas planet atmospheres high temperature kinetics. T6 - Construction of an efficient numerical tool for the reliable simulation of the super-orbital entry conditions T7 - Development of physical-based macroscopic models from accurate state-to-state (STS) models, including the effects of non-equilibrium distributions. T8 - Construction of a database of collision integrals for binary interactions involving the species relevant to a contaminated hydrogen/helium plasmas with impurities (Ar, methane), and species originating from carbon-phenolic ablating TPS T9 - Computation of front and back convective and radiative heat flux on typical probe configurations T10 - Analyse available tube test data (ESTHER, T6 and EAST).
Background and previous activities:
In recent years, the use of state-to-state kinetics as the proper aero-thermodynamic model in conjunction with fluid dynamics conservation laws is more often employed thanks to the availability of increased computational power. As it is well known the common assumption made in multi-temperature approaches of considering a Boltzmann distribution fails because the population of very high energy levels (the most important for chemical processes) are not correlated with the vibrational temperature. Coupling vibrational kinetics with fluid dynamic codes is a complex problem, due to the large dimension of the chemical variables. The number of virtual species is of the order of 100 and the number of processes of the order of 10000. These numbers limit the applicability of the state-to-state model to simple 1D geometries. These problems could be overcome by using high-efficient processing units recently adapted for scientific calculations. New programming tools can be executed efficiently on specialized computers. Today 2D GPU accelerated codes have been already applied to air.
Technical Officer: Louis Walpot ([email protected]) 79 Calibrated Heat flux Device for Thermal Protection System Qualification in Europe’s Aerothermal Testing Facilities Prog: TDE Budget: 600K 18 months Aim: C Ref.: 5 TRL: 4 to 5
Objective(s): Experimental performance evaluation of plasma facilities for Thermal Protection System (TPS)
Description:
The performance of a Thermal Protection System (TPS) is qualified in Plasma Wind Tunnels. The facility performance and flow conditions are typically compared based on cold wall copper heat flux and total pressure.
A major concern for reliable, thorough and accurate cold wall heat flux measurement is the lack of suitably sized sensors dedicated to the harsh environment of high-enthalpy plasma flows with high currents, high voltages and electromagnetic radiation. The surface state of the cold wall is of utmost importance for a quantitative measurement.
It is proposed to develop a sensor system dedicated to these facilities. The development contains not only sensor principle and design, but requires suitable calibration procedure development including surface state definition. The result of this initiative eventually allow the intercomparison of European facilities with respect to the qualification of various flow conditions (pressures, speeds, atmospheres) and eventually flight systems of entry vehicles. Due to the fact that different facilities are used for different trajectory phases, this heat flux sensor system will serve as the basis of the aerothermal characterization of re-entry flight qualification of TPS.
Background and previous activities: Within the ESA study “Standard Model Testing” a common sensor design for European facilities was realized. It was a compromise taking into account the different facilities’ sizes and mechanical constraints. The measured data at comparable flow conditions however was considerably different. An open issue within that study was the surface state and independent calibration of the sensors. This study aims to close the gap.
Technical Officer: Louis Walpot ([email protected]) 80 Computational techniques for fast radiative transfer simulations on distributed memory HPC-architectures. Prog: TDE Budget: 500K 18 months Aim: D Ref.: NEW TRL: 4-5
Objective(s): Implementation, verification and validation of a Solver for the radiative transfer equation which - is applicable to highly parallelized HPC infrastructures with distributed memory - modularly attachable to CFD, DSMC solvers - uses domain decomposition - is applicable to complex geometries - requires a minimum of user interference (e.g. uses automatic generation of radiative sub-meshes, automatic domain decomposition) Description: An accurate computational fluid dynamics analysis of these flows requires adequate modeling of energy transport by radiation. The range of numerical techniques for the treatment of radiation in numerical flow simulations includes Monte Carlo models, which enable accurate and computationally efficient predictions of radiative energy transport in participating (absorbing and emitting) media for complex geometries. A common challenge is the implementation of these methods into highly parallelized high-performance computing environments. Besides the efficient implementation of the radiation solver itself, a comprehensive range of additional problems which need to be addressed includes: • load balancing /weighted partitioning based e.g. on the emission in the volume elements; automatic domain decomposition, especially for statistical particle- based methods • data exchange and accuracy preservation at domain boundaries • Automatic generation of radiative subgrids (due to the high numerical complexity of solutions algorithms of the radiative transfer equations, flow solver grids are likely to be to fine and/or to be resolved according to flow structures which are not primarily relevant to radiation transport. Hence, automatic cell agglomeration, re-meshing or other techniques need to be developed). • Interpolation algorithms of the radiative source term to the flow solver and efficient coupling to radiative property databases such as PARADE. (also for the case of different domain decompositions for the flow- and radiation solver on distributed HPC-architectures.) • Conservative smoothing algorithms to control the noise of statistical methods. • Coupling strategies to DSMC solvers for low-density applications. • A strategy to reduce the spectral complexity to reduce the CPU loads, like e.g. Hybrid Statistical Narrow Band (HSNB)
Background and previous activities: Space exploration missions may include atmospheric entries at high velocities (outer gas planets entries or, sample return or entries from hyperbolic orbits) or into an atmosphere with highly radiative constituents (for example, Mars or Titan). The effects of gas radiation in the shock layer and wake flows may become significant at such flight conditions. Further, flows in plasma-ground test facilities (such as plasmatrons or arc-heaters) are highly affected by radiation. Radiating plumes of solid rocket boosters may cause significant heat loading at the base of launcher configurations.
Technical Officer: Louis Walpot ([email protected]) 81 Development of a dedicated axisymmetric/3D non- equilibrium CFD solver for shock tube facilities Prog: TDE Budget: 500K 24 months Aim: D Ref.: new TRL: 5
Objective(s): Shock-tube experiments are used in space mission design and in particular for the prediction of the spacecraft's radiative heating during Earth re-entry from either a lunar or Mars return mission and outer gas planets entries. Shock tube facilities like ESTHER, T6 in Europe and EAST in NASA Ames, USA have the capability to generate shock-heated test gas at velocities and pressures representative of these return trajectories. Experimental data is used for assessing spacecraft aerothermal loads and the developed tool plays an essential role as a benchmark to the underlying thermochemical nonequilibrium models. The objective of this activity is to develop dedicated high fidelity time-accurate nonequilibrium Navier–Stokes solver to simulate shock tube process for outer gas planet missions on a distributed HPC with the focus on the post-shock region and the absolute radiation emanating from it.
Description: Although 3D time-accurate nonequilibrium codes exists, the activity will focus on developing a dedicated code which can be used to simulate the facilities in reflected mode or unsteady expansion. Several computational challenges and a large number of physical processes have to be addressed: heating process in the driver, diaphragm rupture, turbulent mixing of the driver gas hot jet and the cold driven gas, boundary layer and rarefaction effects, and radiative losses. Furthermore the task is complicated in regards to the times and space scales, as well as chemical and thermal stiffness of the source terms. As such the tool to be developed shall be able to simulate test runs simplifications in the computational modelling and modular inclusion of non-equilibrium effects.
Task: - Design of a computational code based on an already validated time accurate 2D/3D nonequilibrium solver, with efficient parallelization - Introduce models to the heating process in the driver, diaphragm rupture, and radiative losses. - Particular emphasis is placed to reduce the computational times, to track the running shock with adaptive meshing techniques. - Build in levels of simplifications. - validate the solver with experimental data. - Road map for further improvements. - Radiation from the driver effects (the hot driver illuminates the test gas and may in some cases pre-ionize the flow ahead of the shock front).
Background and previous activities: Equilibrium 0D, 1D and 2D simulation tools exists today, like e.g. STAGG (analytical shock-tube code used for predicting the performance of the ESTHER shock-tube) or ANITA and SPARK which are 3D and 2D-axi Navier-Stokes codes with nonequilibrium thermodynamics and chemistry. Although these are developed with the aim of reproducing hypersonic blunt-body flows (except STAGG, developed in the scope of ESTHER development), these have been applied to the simulation of the hydrodynamic properties of shocked flows such as in the ESTHER shock-tube.
Technical Officer: Louis Walpot ([email protected]) 82 Assessment of grid fins designs of first stage recovery
Prog.:TDE Budget: 600K 18 months Aim: D Ref.: New TRL: 2 - 4
Objective(s): The objective of the activity is to design and optimize grid fins for next generation first stage rockets.
Description: Grids fins shapes are used to provide the required torque and forces while minimizing the mass and withstanding structural and thermal loads.
Many possibilities of geometric combination can be analysed to perform shape optimizing of the grid fin during the descent of the first stage rocket. The high heat load can be attributed from the hypersonic phase as well as the interaction of the hot the plume of the retro propulsion system. Additionally the risk of excessive structural loads due to vibrations particularly in the frequency domain that trigger resonance phenomena has to be mitigated. A trade-off shall be carried out considering the refurbishment for re- use.
This activity will encompass the following tasks • Literature survey grid fin designs. • Trajectory analysis for a representative return booster. • Grid fin design optimization based over the complete supersonic till the subsonic regime • Aerodynamic and geometry optimization is to be performed by CFD and experiments to find the optimal combination of parameters providing the best design for the targeted re-flight mission: size, mass, lattice, shape, thermal design (e.g. radiative cooling/ablation). • The design verification and assessment relies on high quality CFD (special regard to results consolidation, grid convergence, turbulence modelling, geometrical details e.g. leading edges), and wind tunnel tests testing large grid enabling to characterise the details of the flow (Schlieren, PSP, Oil flow, PIV). Test should encompass wide flow regimes as well as interference effects from mounting devices. • Derive design guidelines for Space Transportation applications.
Background and previous activities: Grid fin is an aerodynamic control surface that usually used on missiles and rockets. It is also used to control the rocket fly back providing the control during the return to earth from hypersonic to subsonic flows. Recently several study are on going in Europe re-use the first stage rockets like e.g. Ariane Next.
83 Thank you for attending !
Please feel free to contact any of the organisers for additional feedback, remarks or comments:
Csaba Jéger ([email protected]) Stephan Schuster ([email protected]) Orr Cohen ([email protected])
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