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ENGINEERING TOOLS, TECHNIQUES AND TABLESWIND TUNNELS: AERODYNAMICS, MODELS AND EXPERIMENTS JUSTIN D. PEREIRA EDITOR Nova Science Publishers, Inc. New York
CONTENTSPreface vii Chapter 1 Design, Execution and Numerical Rebuilding of Shock Wave Boundary Layer Interaction Experiment in a Plasma Wind Tunnel 1 M. Di Clemente, E. Trifoni, A. Martucci, S. Di Benedetto and M. Marini Chapter 2 The Mainz Vertical Wind Tunnel Facility– A Review of 25 Years of Laboratory Experiments on Cloud Physics and Chemistry 69 Karoline Diehl, Subir K. Mitra, Miklós Szakáll, Nadine von Blohn, Stephan Borrmann and Hans R. Pruppacher Chapter 3 Modeling and Experimental Study of Variation of Droplet Cloud Characteristics in a Low-Speed Horizontal Icing Wind Tunnel 93 László E. Kollár and Masoud Farzaneh Chapter 4 An Air-Conditioned Wind Tunnel Environment for the Study of Mass and Heat Flux Due to Condensation of Humid Air 129 Akhilesh Tiwari, Pascal Lafon, Alain Kondjoyan and Jean-Pierre Fontaine Chapter 5 In-Situ Evaluation for Drag Coefficients of Tree Crowns 147 Akio Koizumi Chapter 6 The Pre-X Lifting Body Computational Fluid Dynamics and Wind Tunnel Test Campaign 167 Paolo Baiocco, Sylvain Guedron, Jean Oswald, Marc Dormieux, Emmanuel Cosson, Jean-Pierre Tribot and Alain Bugeau Chapter 7 Low-Speed Wind Tunnel: Design and Build 189 S. Brusca, R. Lanzafame and M. Messina Index 221
PREFACE This new book presents current research in the study of wind tunnels, including thedesign, execution and numerical rebuilding of a plasma wind tunnel with the aim to analyzeshock wave boundary layer interaction phenomena; the Mainz vertical wind tunnel facilityexperimenting on cloud physics and chemistry; an air-conditioned wind tunnel environmentfor the study of mass and heat flux; using wind tunnel studies to evaluate the drag coefficientof the tree crown and Pre-X aerodynamic/aerothermal characterization through computationalfluid dynamics and wind tunnels. Chapter 1 - The present chapter reports the design, execution and numerical rebuilding ofa plasma wind tunnel experimental campaign with the aim to analyse shock wave boundarylayer interaction phenomena in high enthalpy conditions. This particular flow pattern could arise in proximity of a deflected control surface, thusgenerally causing a separation of the boundary layer and a loss of efficiency of the controlsurface itself; moreover, high mechanical and thermal loads are generally induced at the flowreattachment over the flap. Therefore, the analysis of this problem is crucial for the designand development of the class of hypersonic re-entry vehicles, considering that, even though ithas been widely analyzed in the past, both from an experimental and theoretical point ofview, by describing its physical features, only few studies have been carried to analyse thephenomenon in high enthalpy real gas and reacting flow conditions. The activity has been developed by analysing the flow phenomenon of interest indifferent conditions: i) hypersonic re-entry conditions considering the ESA EXPERT capsuleas a workbench, and ii) ground-based facility conditions considering the CIRA Plasma WindTunnel “Scirocco”. The aim has been the correlation of the results predicted, by means of aCFD code, and then measured through specific experiments suitably designed, in these twodifferent environments. To this effect, a flight experiment has been designed to be flown on the EXPERT capsulealong the re-entry trajectory in order to collect flight data (pressure, temperature and heatflux) on the shock wave boundary layer interaction phenomenon to be used for CFDvalidation and, additionally, as a reference point for the extrapolation-from-flightmethodology developed accordingly. Requirements for the experimental campaign to beperformed in the “Scirocco” facility have been derived considering the most critical andinteresting points along the EXPERT trajectory. A suitable model, representative of theEXPERT geometry in the zone of interest, i.e. the flap region, has been conceived bydefining the main design parameters (nose radius, length, width, flap deflection angle) and an
viii Justin D. Pereiraexperimental campaign has been delineated, the aim being to reproduce on this model thesame mechanical and thermal loads experienced ahead and over the EXPERT full- scale flapduring the re-entry trajectory. Suitable facility operating conditions have been determinedthrough the developed extrapolation-from-flight methodology; the design and the analysis ofshock wave boundary layer interaction phenomenon has been done by focusing the attentionmainly to the catalytic effects over the interaction induced by the different behaviour in termsof recombination coefficient of the materials involved in the problem under investigation. Once defined the design loads, the model has been realized and tested in the PlasmaWind Tunnel Scirocco under the selected conditions. The numerical rebuilding, showing areasonable good level of reproduction, has been also carried out, even though the validationof the entire extrapolation-from-flight and to-flight developed methodology could becompleted only after the EXPERT flight currently planned in mid 2011. Chapter 2 - The Mainz vertical wind tunnel is so far a worldwide unique facility toinvestigate cloud and precipitation elements under conditions close to the real atmosphere.Hydrometeors such as water drops, ice crystals, snow flakes, and graupels are freelysuspended at their terminal velocities in a vertical air stream under controlled conditionsregarding temperature (between -30°C and +30°C), humidity (up to the level of watersaturation), and laminarity (with a residual turbulence level below 0.5%) of the air stream.Cloud processes in warm, cold, and mixed phase clouds have been investigated in the fieldsof cloud physics and chemistry, aerosol–cloud interactions, and the influence of turbulence.The experiments include the behaviour of cloud and rain drops, ice and snow crystals, snowflakes, graupel grains and hail stones and the simulation of basic cloud processes such ascollisional growth, scavenging, heterogeneous drop freezing, riming, and drop-to-particleconversion. Atmospheric processes have been investigated under both laminar and turbulentconditions in order to understand and quantify the influence of turbulence. The results are essential for applications in cloud chemistry models to estimate theatmospheric pathway of trace gases, in cloud and precipitation models to improve thedescription of the formation of precipitation (growth and melting rates), and in now- andforecasting of precipitation to improve the evaluation of radar and satellite data. Chapter 3 - Variation of the characteristics of aerosol clouds created in icing wind tunnelsis studied theoretically and experimentally. The characteristics of interest are the droplet sizedistribution, liquid water content, temperature, velocity, and air humidity, which are amongthe most important factors affecting atmospheric icing. Several processes influence thetrajectory, velocity, size and temperature of the droplets, such as collision, evaporation andcooling, gravitational settling, and turbulent dispersion. The authors have developed a two-dimensional theoretical model that takes these processes into account, and predicts how theyinfluence the changes in the characteristics of the droplet cloud during its movement in thetunnel. The most recent development pays special attention to two of the possible collisionoutcomes, i.e. coalescence after minor deformation and bounce, together with the transitionbetween them. Indeed, these outcomes are frequent when the relative velocity of the dropletsis small, as is the case for a cloud formed after the injection of water droplets in the directionof air flow. An experimental study is also carried out with different thermodynamicparameters at different positions in the test section of the tunnel, which makes it possible toobserve the evolution of cloud characteristics under different ambient conditions. The dropletsize distribution and liquid water content of the aerosol clouds were measured using anintegrated system for icing studies, which comprises two probes for droplet size
Preface ixmeasurements and a hotwire liquid water content sensor. Droplet trajectories were observedusing particle image velocimetry. The experimental results are also used to validate the modelby comparing them to model predictions. Satisfactory agreement between the experimentaland calculated results establishes the applicability of the model to determine the evolution ofdroplet size distribution and liquid water content in an aerosol cloud in the streamwisedirection, together with their vertical variation. Chapter 4 - The development of an artificial ecosystem inside a closed environment isone of the future challenging problems, which is mandatory for the long duration mannedspace missions like lunar base or mission to Mars. Plants will be essential companion lifeforms for such space missions, where human habitats must mimic the cycles of life on earthto generate and recycle food, oxygen and water. Thus the optimized growth of higher plantsinside the closed environment is required to obtain efficient biological life support systems.The stability and success of such systems lie on the control of the hydrodynamics and on anaccurate characterisation of the coupled heat and mass transfer that develop at interfaces(solids, plants,..) within the space habitat. However, very few data can be found on the precisecharacterization / prediction of the mass transfer at interfaces, and more particularly in space.In most studies the mass flux is deduced from the measured / calculated heat flux by a heatand mass transfer analogy. Hence, the authors have developed a ground based experimental set-up to measure the airflow velocities and concomitant mass transfer on specific geometries under controlled airflow conditions (flow regime, hygrometry, temperature). The final goal is to derive atheoretical model that could help for the prediction of the hydrodynamics and coupledheat/mass transfer on earth, and eventually in reduced gravity. The authors have used aclosed-circuit wind tunnel for our experiments, which can produce very laminar to turbulentflows with controlled temperature and hygrometric parameters inside the test cell. The initialexperiments have been performed in dry air with an average velocity between 0.5-2.5 m.s-1.The velocity profiles near a clean aluminium flat plate in horizontal or vertical positions havebeen studied for low Reynolds number flows by hot wire anemometry. The measurementswith the horizontal plate showed a boundary layer thickness in agreement with the Blasius’solutions. Condensation of humid air was induced on an isothermal flat plate, which wascooled by thermoelectricity. The mass transfer on the plate was controlled and recorded witha precise balance. The obtained results are analyzed, and compared to the available data oncondensation. Chapter 5 - In order to make a hazard prediction of trees against wind damage, such asstem breakage or uprooting, it is essential to quantitatively estimate the wind force acting on atree. The drag coefficient of the tree crown, which is necessary to estimate wind force, hasbeen evaluated using wind tunnel studies. Most of the specimens used for wind tunnel studieswere dwarf trees, because of the restrictions due to wind tunnel size. However, with regard tothe wind-force response, the similarity rule is not applicable to the relationship between dwarftrees and actual-sized trees. In fact, the drag coefficients of small trees were found to beconsiderably greater than those of actual-sized trees. To estimate the drag coefficients ofactual-sized trees accurately and easily, a field test method was developed. Using this method,wind speed and stem deflection were monitored simultaneously. The wind force acting on thetree crown was calculated from the stem deflection; the stem stiffness was evaluated byconducting tree-bending tests. The field tests were conducted on black poplars and a Norwaymaple; the results showed that the drag coefficients decreased with an increase in wind speed.
x Justin D. PereiraThis decrease can be explained mainly by the decrease in the projected area of the crown,because of the swaying movement of the leaves and branches. Although the variation in thedrag coefficients was large at low wind speeds because of the swaying behavior of the stemsubjected to a variable wind force, the variation at wind speeds above 10 m/s was small. Theaverage drag coefficient for black poplars at a wind speed of 30 m/s was estimated by thecurve fitting of a power function to the wind velocity-drag coefficient relationship, and thisvalue was found to be not greater than that of actual-sized conifers previously studied in windtunnel experiments. These results suggest that the wind permeability of poplar crowns isgreater than that of conifer crowns due to the difference in leaf flexibility. Although the dragcoefficients in the defoliation season were smaller than those measured in the leaved season atlow wind speeds, the difference in drag coefficients became less pronounced at high windspeeds. Chapter 6 - Pre-X was the CNES proposal for demonstrating the maturity of Europeantechnology for gliding re-entry spacecraft. The program finished in year 2007 with the end ofthe phase B and a successful PDR. Then it was stopped with the aim of joining the ESAproject IXV. The main goal of this experience is to demonstrate the implementation of reusablethermal protections, perform aero thermo dynamics experiments and efficiency of a suitableguidance navigation and control system. The attitude control is realised by elevons andreaction thrusters overall the hypersonic flight, with a functional and experimental objective. This paper presents the Pre-X aerodynamic / aerothermal characterisation throughcomputational fluid dynamics and wind tunnels tests performed during the phases A and B ofthe programme. The tests permitted to cover the Mach range from 0.8 to 14 and to investigatethe main effects of aerodynamic and aerothermal phenomena. In the preceding phases theaerodynamic shape and centring had been defined. The logic and main results of this activity are presented in this paper. Chapter 7 - In this chapter the authors deal with a procedure for the design and build of alow speed wind tunnel for airfoil aerodynamic analyses and micro wind turbine studies. The designed closed-circuit wind tunnel has a test chamber with a square cross section(500 mm x 500 mm) with a design average flow velocity of about 30 m/s along its axis. The designed wind tunnel has a square test chamber, two diffusers (one adjacent to thetest section and one adjacent to the fan to slow the flow), four corners (with turning vanes) toguide the flow around the 90° corners, an axial fan to guarantee the mass flow rate andbalance any pressure loss throughout the circuit, a settling chamber with a honeycomb (toeliminate any transverse flow), a series of ever-finer mesh screens (to reduce turbulence) anda nozzle to accelerate flow and provide constant velocity over the whole test chamber. Thepressure losses of single components were evaluated as well as the global pressure loss (thesum of pressure losses of all the single components). Once the pressure losses were evaluated,the axial fan was chosen to guarantee the design’s volumetric flow, balance pressure lossesand above all maximise its performance. The definitive dimensions of the wind tunnel are10.49 m x 3.65 m. Once the design targets were defined, the test chamber dimensions, maximum wind speedand Reynolds numbers were calculated. At the end of the design process, the wind tunnel energy consumption was estimated andon-design and off-design performance was evaluated to obtain the wind tunnel circuitcharacteristics for a defined velocity range (0 – 50 m/s).
Preface xi The best circuit and axial fan matches were performed in both the open and closed testsection configurations. Using the matching procedure between the fan and wind tunnel’smechanical characteristics (global pressure loss as a function of wind velocity), the fanoperating parameters were set up for optimum energy conservation.
2 M. Di Clemente, E. Trifoni, A. Martucci et al. and heat flux) on the shock wave boundary layer interaction phenomenon to be used for CFD validation and, additionally, as a reference point for the extrapolation-from-flight methodology developed accordingly. Requirements for the experimental campaign to be performed in the “Scirocco” facility have been derived considering the most critical and interesting points along the EXPERT trajectory. A suitable model, representative of the EXPERT geometry in the zone of interest, i.e. the flap region, has been conceived by defining the main design parameters (nose radius, length, width, flap deflection angle) and an experimental campaign has been delineated, the aim being to reproduce on this model the same mechanical and thermal loads experienced ahead and over the EXPERT full- scale flap during the re-entry trajectory. Suitable facility operating conditions have been determined through the developed extrapolation-from-flight methodology; the design and the analysis of shock wave boundary layer interaction phenomenon has been done by focusing the attention mainly to the catalytic effects over the interaction induced by the different behaviour in terms of recombination coefficient of the materials involved in the problem under investigation. Once defined the design loads, the model has been realized and tested in the Plasma Wind Tunnel Scirocco under the selected conditions. The numerical rebuilding, showing a reasonable good level of reproduction, has been also carried out, even though the validation of the entire extrapolation-from-flight and to-flight developed methodology could be completed only after the EXPERT flight currently planned in mid 2011. 1. INTRODUCTION The high cost of access to space is the main limitation to scientific research and spacecommercialization, and for this reason all the countries in Europe are thinking how designadvanced spacecrafts in order to achieve low launch costs in the near future (Ref. ,).Spacecrafts like the US Space Shuttle Orbiter represent the first generation of reusable launchsystems but several system studies have been conducted during the 80’s to investigatepossible future concepts for the next generation of RLVs. In the frame of the ESA-FESTIPProgram in late 90’s, system concept studies were carried out and an extensive investigationof a wide range of RLV concepts (more than 10 configurations) was performed (ref. ). Inthe following decade several programmes, at European and national level, were launched topromote the development of some of the identified enabling technologies required for thefuture generation of reusable space transportation systems that shall be safer and lessexpensive with respect to the US Space Shuttle. Enabling technologies for such vehicles andderived systems must be inherently reliable, functionally redundant, wherever practical anddesigned to minimize or eliminate catastrophic failure modes. Reliability could be improvedthrough performance margin that translates to robust design, and this presupposes thematuration of some specific macro-technologies: • Re-entry heating. the aerospace vehicles have to handle the typical large thermal loads encountered during re-entry to Earth from LEO, due to the necessity of reducing the vehicle speed before landing; • Hypersonic flight navigation. the future space vehicles will have to fly for large part of their mission to speed much greater that the speed of the sound, and will have to maneuver safely in such conditions;
Design, Execution and Numerical Rebuilding of Shock Wave… 3 • Reusability. the most important characteristic from the operational point of view is the tendency to be as much like current airplanes thus translating into the reusability concept. Starting from the necessity of a proper level of maturity of these high level technologies,some guidelines and critical points to be developed at lower level, in order to match theparticular requirements for the RLV design, were identified by past space systems andtechnological programs. Among the others, can be identified: • Configuration Design • Extrapolation to Flight • Transition Prediction • Control Surface Aerothermodynamics It is clear that many other technological areas are being involved and ask for othersignificant developments (propulsion, flight mechanics, stability and control, guide andnavigation, configuration optimization, etc.) but, in any case, to develop the future spacetransportation system a considerable work should be devoted to the aeroshape definition inorder to improve performance, flyability and controllability, propulsion integration, heat loadreduction, stage separation, coupling between forebody aeroelasticity and propulsion system,coupling between viscous drag and heat loads. The aerodynamic efficiency (E=CL/CD) shouldbe increased since they will experiment large part of flight at moderate altitude at high Machnumber, strongly asking for more efficient aerodynamic design. Also the transition processfrom laminar to turbulent boundary layer should be predicted with greater accuracy since itplays an important role in the design of aerospace planes thermal protection system, and thecurrently available theoretical know-how (i.e. the stability theory) could not yet guarantee fora safe and reliable transition prediction (Ref. , , ). Among the others, the study of aerodynamic efficiency of control surfaces plays a role ofprimary importance (Ref. ). In fact, the necessity of manoeuvrability and high cross-rangeduring ascent or re-entry phase requires the capacity to increase control surfaces aerodynamicefficiency whose analysis is strictly connected to the study of shock wave boundary layerinteraction (SWBLI) occurring around them. The increase of knowledge must regard,especially in the SWBLI phenomenon in high enthalpy conditions, the prediction, with a goodlevel of approximation, of its behaviour in flight conditions. In a classical approach, the design of space vehicles (e.g. the Space Shuttle) is basedheavily upon experimental data although, due to the inherent limitations of similarity laws,ground based facilities cannot simulate completely the physics of flows experienced by suchvehicles during re-entry. To overcome these limitations different strategies could be adopted:in US data obtained from in-flight experiments, particularly with the X-series vehicles, havebeen used to complement the test data obtained from ground-based facilities; on the otherhand, since the times of Hermes Program, Europe chose to complement the knowledgeavailable from the cold wind tunnels, which are not able to model the high-temperature andreal gas effects typical for higher speeds and altitudes, by means of high enthalpy or hot-flowfacilities. ESA, therefore, supported the updated of existing cold flow wind tunnels, and alsothe construction of facilities with new capabilities, as for example the PWT Scirocco of
4 M. Di Clemente, E. Trifoni, A. Martucci et al.CIRA, to investigate the heat loads and gas surface interactions on materials and large sizestructures (ref. , ). In any case, the prediction of hypersonic flows, both for the complexity of the requiredphysical modelling and for the impossibility to duplicate in wind tunnels real flight conditionsdue to the high energy required, is still one of the main problem related to the development ofthe new class of space vehicles. Moreover, high efficiency space vehicles require complexinvestigations because of the large contribution of the viscous effects to the aerodynamicforces and heating, while the effects of the gas modelling are important since the small bluntnose, necessary to increase the aerodynamic efficiency, does not shield the rest of the vehicle,thus implying the presence of large chemical effects on most of the vehicle surface. The main data sources for the aerothermodynamic design of a space vehicle arecomputational fluid dynamics, wind tunnel tests and flight experiments, generally onsimplified geometries: • wind tunnel tests are important because they allow carrying on “controlled” simulations and therefore to better understanding the flow-physics phenomena; although ground-based facilities provide fundamental information for flight, no one facility can provide all of the aerothermodynamic information required for the design of a vehicle. As today it is well recognised, duplication of all flight characteristic parameters (Mach, Reynolds, Damkhöler, state of the gas) in a ground facility is not possible, particularly flight Reynolds number and high enthalpy effects are critical and difficult to be reproduced at ground; • flight experiments data represent the “truth” to be predicted, i.e. they show the real performance of the vehicle in representative conditions and, therefore, they are unique for vehicle qualification although they are quite costly, require considerable time and have uncertain repeatability and accuracy. Many phenomena can not be directly measured and de-coupling of effects is not always an easy task; • numerical simulations still play a fundamental role in the study of aerot- hermodynamics; moreover, the highest confidence in any ground-based or flight data set occurs when the results obtained with CFD are in agreement with them, the so- called extrapolation-to-flight technique. Even if today CFD is contributing significantly to the aerothermodynamic design of advanced vehicles, it still suffers from lack of physical modelling, robustness and accuracy of the mathematical algorithms, grid generation flexibility and hardware limitations; thus good wind- tunnel and flight data are still necessary for validation and/or calibration of CFD codes used to predict surface and flow field variables for the full-scale vehicle at re- entry flight conditions. The best approach for improving confidence in aerothermodynamic design tools, from acomputational and ground-based experimental point of view, is to validate those tools anddesign approaches with respect to flight experiments. As matter of fact, although in the lastyears Europe has dedicated significant effort to improve the quality and reliability ofaerothermodynamic predictions, due to their key importance in the design and developmentof any hypersonic space vehicle, and a considerable effort has been devoted to the realizationof ground based plasma facilities and development of advanced numerical tools with the state
Design, Execution and Numerical Rebuilding of Shock Wave… 5of the art physical model, in-flight experimentation is still needed to validate thecomputational codes and to establish meaningful and reliable ground-to-flight extrapolationmethodology. Above Mach 10, where in particular high-temperature effects become dominant, CFDrepresents the only prediction tool, and therefore the appropriate validation of numericalcodes is a great concern. Generally it is achieved, by comparing data measured in high-temperature facilities with those obtained by numerical prediction and in many cases anumerical approach is used to define the experimental test cases and for the interpretation ofthe measured data. CFD codes are subsequently being used for flight simulations above Mach10 even if this ‘extrapolation method’ assumes that the physical models enabling good resultsfor the simulation of the experimental test cases, provide good results also for free flight.Therefore, free-flight data are required to remove any doubt about the validity and accuracyof the CFD predictions, and to confirm the extrapolation methodology as well. The mainargument for the in-flight experimentation is therefore the need for realistic and combinedloads levels which are representative for the operational environment of a RLV. Such testsmust be performed complementarily to on-ground testing for validating critical enablingtechnologies of the reference RLV concept. The analysis of this phenomenology is complicated by the fact that, in hypersonic regime,scaling laws have not yet been found. Plasma Wind Tunnels, which allow the same energylevels of the real flight, are in fact characterized by the test chamber flow rather dissociatedconditions, and this has a large influence on the flow-field around the test article, while coldhypersonic wind tunnels, where the simulation is focused on the duplication of Mach andReynolds numbers, permit only to reproduce the classical aerodynamic forces and the relatedcoefficients even though with strong limitations. The influence of real gas effects and viscousinteraction effects on control flap efficiency and heating is one of the mainaerothermodynamic issues for the next generation RLV design, together with the qualificationtesting of the thermal protection system in ground-based facility and the consequentextrapolation to flight for experimental results. In order to assess these issues, a numerical approach has been followed to define a wind-tunnel experimental campaign on a representative model to reproduce the in-flight expectedvalues of mechanical and thermal loads acting on a typical control device, in interpreting themeasured data and finally for the extrapolation to the flight conditions of the experimentalresults, as the local conditions in the wind tunnel facility only partially duplicate those inflight. Moreover, a flight experiment whose results could be used as point of reference forsuch phenomena has been also designed. As matter of fact, aerothermodynamic design issues,as the analysis of flap efficiency for control and navigation, has been addressed usingadvanced numerical codes, ground-based facilities and flight testing. Following the previous considerations, it arises the need to develop an extrapolation-from-flight and to-flight methodology able to combine and mutually validate the flight andground data on the problem of interest. Even though the prediction of mechanical and thermalloads acting on the control surfaces of hypersonic vehicles is crucial for the design of theiraerodynamic shapes and thermal protection system, at the moment the lack of hypersonicflight data that can serve as a point of reference for the validation process, makes itimpossible, especially for some of the most challenging hypersonic problems.
6 M. Di Clemente, E. Trifoni, A. Martucci et al.Figure 1. Extrapolation from flight and to flight procedure. The issue of an extrapolation-to-flight methodology for high enthalpy flow must be in thelight of a progressive building up of confidence in the design of a space vehicle. Thedevelopment of such methodology, whose rationale is shown in Figure 1, has been carried outreferring to the ESA EXPERT capsule (ref. ), which has the indubitable advantage to be asimple geometry, conceived as an experimental test-bed for in-flight experimentation anddesigned to avoid degradation and flow contamination. In order to develop the methodologyand to extrapolate plasma facility results to real flight conditions, it is necessary, first of all, tocharacterize by means of CFD simulations the flight conditions in the flap region and design aflight experiment in order to instrument the vehicle and to collect flight data during the re-entry mission, to be used for the post-flight analysis to validate the entire procedure. On theother hand it is necessary to design, perform and numerically rebuild a number ofexperimental tests in a plasma wind tunnel facility that can be representative of the flightconditions with respect to the SWBLI phenomenon over the flap. Finally, plasma wind tunnelresults must be correlated, by means of the relevant parameters of the interaction as viscousinteraction parameter and rarefaction parameter, with those predicted (and then measured)during the flight, the goal being to understand the test conditions necessary to reproduce(simultaneously or separately) the mechanical (pressure) and thermal (heat flux, temperature)loads acting on the control surface device.1.1. EXPERT Capsule The development of the extrapolation to flight methodology has been carried outreferring to the ESA EXPERT capsule whose in-flight test program focuses on a genericcapsule-like configuration designed in such a way to enhance the most interestingaerothermodynamic phenomena of a typical re-entry vehicle performing a sub-orbital ballistichypersonic flight. The main objective of the project is to collect in-flight data on the mostcritical aerothermodynamic phenomena via dedicated classical and advanced flight testmeasurement assemblies (i.e. EXPERT Scientific Payloads), and this in order to improve theknowledge about the differences between ground experiments and real flight conditions; eachparticular phenomenon related with the high energy re-entry mission (gas-surface interaction,induced and natural laminar-to-turbulence transition, real gas effects on shock wave boundary
Design, Execu D ution and Num merical Rebuilding of Shoc Wave… ck 7la ayer interactio shock lay chemistry has been separately an on, yer y) nalyzed, and a specific xperiment has been design for each of them in t frame of the Technica Researchex s ned the alPrrogram related to the capsu developme The scient d ule ent. tific data will then be used to validatest tate-of-the-art numerical to ools for aeroothermodynam application and groun mic ns nd-to-flightex xtrapolation pr rocedures (Re ). ef. Each Paylo will be q oad qualified acco ording to the relevant Ass e sembly, Integgration andVerification Plan; in paraV P allel a numb ber of exper rimental activ vities in the field of eae erothermodynamics are carr out to acq ried quire all necessary pre-flight information on specificph henomena allo owing for an optimized pos st-flight phase Among the others, specia attention e. alha been given to the Shock Wave / Boun as n k ndary Layer In nteraction phe enomenon wh hose effectson the open flap are being in n ps nvestigated w two differe Scientific Payloads, i.e. Payload 6, with entth hrough instrummentation of f flaps and cavi ities with mai inly classical sensors, and Payload 7,th hrough the cha aracterization of the bounda layer appr ary roaching the f flap, whose deevelopmentha been carried out in the fra of the pre as d ame esent research activity. The referen geometry of the EXPER capsule is a body of rev nce RT s volution with an ellipse-cl lothoid-cone two-dimension longitudina profile cut b 4 planes an equipped w 4 fixed nal al by nd withop flaps. Th elliptical n pen he nose has a ra adius of 600 mm at the s stagnation point and anec ccentricity of 2.5. The fixed flaps have a deflection of 20 deg, a wid of 400 mm and an x- d f dth max projected length of 300 m (see Figur 2). xis l mm reFi igure 2. EXPER capsule. RT 2. 2 MATHEM MATICAL MODEL The analys of shock wave bounda layer inte sis ary eraction, for t developm the ment of theex xtrapolation-fr rom-flight me ethodology, ha been carried out consider as d ring the CIRA numerical Aco ode H3NS which allows for the ae w s erothermodyna amic analysis over comp s plex three-di imensional ge eometries and suitable to s d simulate commpressible flow at high ent w thalpy (ref.[2 28]). One of the main charact personic flows is that, due to the high temperatures teristics of hypex xperienced be ehind the bow shock, the gas cannot be considere as a perf w e ed fect gas asco ommonly assu umed for low speed flows; air molecule at temperat w ; es tures higher th 800 K hanst to vibrate and for temp tart e peratures arou 1500 K t dissociatio of oxygen molecules und the onbe egin whereas for higher tem mperature also nitrogen mo olecules disso ociate. The mo odelling ofhi igh temperatu phenome ure ena is quite difficult bec cause of the difference a among the
8 M. Di Clemente, E. Trifoni, A. Martucci et al.characteristic time of fluid-dynamics and that of chemical reactions and vibration. Thissituation generally leads to the thermochemical non-equilibrium. In fact, there are manyproblems in high-speed gas dynamics where the gas doesn’t reach the equilibrium state; atypical example is the flow across a shock wave, where the pressure and temperature arerapidly increased within the shock front. By the time equilibrium properties have beenapproached, the fluid element has moved a certain distance downstream of the shock front.The modelling of these phenomena cannot be limited to a calculation of equilibriumconditions at a certain temperature and pressure, but a number of equations must be added tothe classical Navier-Stokes formulation, one for each vibrating or dissociated species. In the code considered for numerical computations, governing equations have beendiscretized using a finite volume technique with a centred formulation over structured multi-block meshes. This approach is particularly suitable to the integral form of the equations; infact, in a first order approximation, it is simply obtained by integrating the equations for eachcell and considering the variables constant inside each volume. The integral formulationensures that mass, momentum and energy are conserved at the discrete level. Suitable modelshave to be taken into account to define the thermodynamics, the transport coefficients andturbulent variables as reported in detail in Appendix 1. The inviscid fluxes at cell interfacesare computed using a Finite Difference Splitting (FDS) Riemann solver, which is especiallysuitable for high speed problems (Ref.). This method solves for every mesh interval theone-dimensional Riemann problem for discontinuous neighboring states (the states at bothsides of the cell face). The second order approximation for FDS is obtained by means of ahigher order ENO (Essentially Non Oscillatory) reconstruction of interface values. Theviscous fluxes are calculated by central differencing, i.e. computing the gradients of flowvariables at cell interfaces by means of Gauss theorem. Time integration is performed byusing an Euler forward scheme with a semi-implicit pre conditioner based on the derivative ofthe source chemical and vibrational terms. 3. PWT SCIROCCO EXPERIMENT PRELIMINARY DESIGN A number of experiments to be performed in the CIRA Plasma Wind Tunnel “Scirocco”,representative of the capsule flight conditions with respect to the shock wave boundary layerinteraction phenomenon occurring around the 20 deg flap, has been designed: PWT drivingconditions, model configuration and attitude and model instrumentation have been defined,by means of a massive CFD activity performed by using the CIRA code H3NS. Theseexperiments have been designed in order to allow for the duplication of characteristicparameters (viscous interaction parameter, rarefaction parameter, reference pressure and heatflux) of the interaction to reproduce on a full-scale flap model both pressure and heat fluxlevels estimated in critical re-entry flight conditions. The final goal has been to develop anextrapolation-to-flight methodology for such flows since the full duplication of flowcharacteristic numbers (Mach, Reynolds, Damkhöler) and state of the gas is not feasible inground facilities. A parametric analysis of the facility operating conditions and model characteristicdimensions (nose radius, length, flap dimensions, etc.) has been carried out in order to definethe operating conditions and experimental set up that permit a simultaneous reproduction of
Design, Execution and Numerical Rebuilding of Shock Wave… 9mechanical and thermal loads acting on flap in flight conditions over the selected model. Therest of the activity has been devoted mainly to the choice of the different protection materials,its equipment (sensors distribution) and the experimental tests at the selected flow conditions. The final model configuration reproduces the full-scale EXPERT 20 deg flap, mountedon a holder composed by a flat plate with a rounded leading edge (made of copper andactively cooled) and lateral edges. The flap has been realized in C-SiC, which is the samematerial foreseen for the realization of the EXPERT capsule flap, whereas for the flat plate ithas been used Haynes 25 whose main thermo-mechanical characteristics are quite similar toPM1000 which is the material foreseen on the capsule. The effects of catalysis jump, due tothe coupling of different materials, have been analysed by considering the availablerecombination coefficients for the materials of interest, or modelling the different parts asfully or not catalytic.3.1. PWT “Scirocco” Facility Description The CIRA Plasma Wind Tunnel “Scirocco” is devoted to aerothermodynamic tests oncomponents of aerospace vehicles; its primary mission is to simulate (in full scale) thethermo-fluid-dynamic conditions suffered by the Thermal Protection System (TPS) of spacevehicles re-entering the Earth atmosphere. “Scirocco” is a very large size facility, whose hypersonic jet has a diameter size up to 2 mwhen impacts the test article and reaches Mach number values up to 11. The jet is thencollected by a long diffuser (50 m) and cooled by an heat exchanger. Seventy MW electricalpower is used to heat the compressed air that expands along a convergent-divergent conicalnozzle. Four different nozzle exit diameters are available: 0.9, 1.15, 1.35 and 1.95 m,respectively named C, D, E and F. The overall performance of “Scirocco” in terms ofreservoir conditions is the following: total pressure (P0) varies from 1 to 17 bar and totalenthalpy (H0) varies from 2.5 to 45 MJ/kg. Facility theoretical performance map in terms ofreservoir conditions produced by the arc heater is shown in Figure 3. Lower enthalpy valuesare obtained by using a plenum chamber between the arc heater column exit and the nozzleinlet convergent part, which allows transverse injection of high pressure ambient air to reducethe flow total enthalpy.Figure 3. Arc heater theoretical performance map.
10 M. Di Clemente, E. Trifoni, A. Martucci et al. The energetic heart of the facility is the segmented constricted arc heater, a column with amaximum length of 5.5 m and a bore diameter of 0.11 m. At the extremities of this columnthere are the cathode and the anode between which the electrical arc is generated. A powersupply feeds the electrical DC power to the electrodes for the discharge. A compressed airsupply distributes dry compressed air to the various segments of the arc heater column, beingable to supply a mass flow rate ranging from 0.1 to 3.5 kg/s, heated up to 10000 K. The last important subsystem of “Scirocco” is the vacuum system, which generates thevacuum conditions in test chamber required by each test. The system consists of ejectors thatmake use of high pressure water steam as motor fluid (30 bar and 250 °C). The achievementof the operating conditions (P0, H0) in test chamber is assured by the presence, before theinsertion of the model, of a 100mm-diameter hemi-spherical calibration probe made ofcopper, cooled, that measures radial profiles of stagnation pressure (Ps) and stagnation heatflux (Qs) at a section 0.375 m downstream of the conical nozzle exit section, by means ofhigh precision pressure transducers and Gardon-Gage heat flux sensors, respectively. Facilityregulations (mass flow, current) are tuned in order to measure on the calibration probe acertain couple of values (Ps, Qs) which correspond to the desired set point in terms of thecouple (P0, H0).3.2. Facility Performance Evaluation The definition of a representative experiment in the CIRA Plasma Wind Tunnel“Scirocco” has been done by considering the most interesting points of the EXPERTreference trajectory, i.e. point P1, that is the point characterized by the highest stagnationpoint heat flux, and point P2, characterized by high heat flux and a relatively low pressure,potentially critical for passive/active oxidation transition of the C-SiC, which is the materialof the nose and the flaps of the capsule. A preliminary analysis of PWT Scirocco capabilities for the duplication of SWBLI flowshas been carried out, the aim being to understand what it is possible to reproduce in thisplasma facility in terms of the characteristic parameters of the interaction as pressure and heatflux (peak values and reference values, upstream of the separation), viscous interactionparameter, (χ L ) ( ) ≈ M ∞ / Re ∞L , rarefaction parameter, VL ≈ M ∞ / Re ∞L , separation 3length experienced during the flight that has been preliminary predicted through CFDsimulations. Even if in this phase preliminary flight values have been used, it was importantonly to develop an extrapolation from flight methodology that allowed to set the facilityoperating parameters necessary to duplicate assigned flight values, that have been thenupdated through three-dimensional non equilibrium computations. Starting from the nominal operating envelope of PWT facility, considering the conicalnozzle D (length 3.1 m from the throat section, exit diameter 1.15 m) and assuming fullylaminar flows, a certain number of numerical simulations has been performed, basing on thecurrently explored region of the envelope, where qualification and validation tests have beenalready executed in order to have a clear idea of what can be simulated in terms of SWBLIinteresting parameters inside the facility; in particular, the effects of total pressure (at low,medium and high total enthalpy) and total enthalpy (at low, medium and high total pressure)have been investigated. Then, additional computations have been performed in order to
Design, Execution and Numerical Rebuilding of Shock Wave… 11duplicate the estimated flight values of the interesting parameters of the interaction around thebody-flap. The complete CFD test matrix, for this preliminary phase, is reported in Table 1 interms of total pressure (P0) and total enthalpy (H0), while in Figure 4 these points have beenshown inside the PWT Scirocco theoretical envelope. Table 1. CFD Matrix for preliminary computations P0 (bar) H0 (MJ/kg) PWT-1 2.45 11.90 PWT-2 2.45 15.00 PWT-3 2.40 18.80 PWT-4 4.70 10.40 PWT-5 5.20 15.00 PWT-6 5.20 18.80 PWT-72 7.90 11.00 PWT-8 7.90 15.00 PWT-9 7.90 18.80 PWT-10 13.00 11.00 PWT-111 12.00 15.00 PWT-121 10.00 15.00 PWT-13 7.90 17.90 PWT-14 10.00 11.00 45 40 35 PWT Operating Envelope Preliminary Computations 30 Additional Computations H0, MJ/kg 25 20 15 10 5 0 0 2 4 6 8 10 12 14 16 18 P0, barFigure 4. PWT envelope and operating conditions. For each of the considered points, the PWT nozzle flow has been simulated in thehypothesis of fully laminar thermo-chemical non equilibrium flow, then the centrelineconditions at X = Xnozzle exit + 0.15m (= 3.25 m) have been taken as free stream conditions toperform the simulation of flow around the model (located preliminarily 0.15 m downstreamof the nozzle exit section).2 For these conditions an angle of attack of the model equal to 10 deg has been also considered
12 M. Di Clemente, E. Trifoni, A. Martucci et al. This CFD-based procedure was also successfully applied to previous experiments design,where it was clearly shown that the same test chamber flow is predicted by means of theuncoupled (single simulation of nozzle flow with geometrical extension to the modelstagnation point section) as done in this work and coupled (complete simulation of flowthrough PWT facility, with test chamber details) simulations, thus assessing the accuracy ofthe overall experimental test design. The preliminary geometry, which can be considered representative of the EXPERTgeometry around the body-flap region (scale 1:2), was a 0.60 m long blunted flat plate with a0.15 m long flap forming a 20 deg angle with the plane and a cylindrical nose whose radiuswas equal to 0.25 m as schematically reported in Figure 5.Figure 5. Preliminary shape of the test article. All the computations have been performed in non-equilibrium fully laminar conditions,assuming a fully catalytic wall with a fixed temperature equal to 300 K or the radiativeequilibrium condition, considering the symmetry plane of the model with a two dimensionalapproach. For numerical reasons a horizontal plate has been considered at the flap trailingedge; the effect of this plate, that could fix the reattachment at the end of the flap, has beenanalyzed for the conditions PWT-7 and PWT-3 of Table 1, that are the conditions,respectively, characterized by the highest and the lowest Reynolds number. This effect seems to be negligible (see Figure 6 and Figure 7) being the reattachmentmechanism not “driven” by the expansion at the flap trailing edge, but it occurs on the flap“far enough” from its trailing edge. From the computed results of the present analysis, andfrom considerations about the complexity of baseflow (useless) prediction, it can beconcluded that the geometry model with the flat plate extension behind the flap can be alwaysemployed for present simulations. mach 7.5 0.4 7 1.4 6.5 6 0.3 5.5 Y (m) 1.2 5 4.5 0.2 4 3.5 1 3 0.1 2.5 2 0.5 0.6 0.7 0.8 X (m) 0.9 1 0.8 1.5 Y (m) 1 0.5 0 0.6 -0.5 0.4 0.2 0 0 0.2 0.4 0.6 0.8 1 1.2 1.4 X (m)Figure 6. Mach contours for the condition PWT-7 with base flow.
Design, Execution and Numerical Rebuilding of Shock Wave… 13 1200 200000 1200 200000 Pressure 180000 180000 Pressure - Base 1000 Heat Flux 1000 Heat Flux - Base 160000 Pressure 160000 Pressure - Base Heat Flux 140000 Heat Flux - Base 140000 Heat Flux (W/m2) Heat Flux (W/m2) 800 800 Pressure (Pa) Pressure (Pa) 120000 120000 600 100000 600 100000 80000 80000 400 400 60000 60000 40000 40000 200 200 20000 20000 0 0 0 0 0.4 0.6 0.8 0.4 0.6 0.8 X (m) X (m)Figure 7. Pressure and Heat Flux distribution for PWT-7 (left) and PWT-3 (right). The pressure distribution along the wall for all the computations is reported in Figure 8(left); after the rapid expansion starting from the stagnation point, a quasi-constant pressureregion is observed along the flat plate up to the zone of shock wave boundary layerinteraction induced by the presence of the flap; at the separation location there is a pressurejump due to the separation shock, then pressure reaches a plateau in the recirculation regionand a peak after the reattachment on the flap followed by a sharp expansion at the end of theflap. The overall distribution is typical of such SWBLI interaction. Pressure levels on themodel are mainly influenced by the value of the total pressure being negligible the effect ofthe total enthalpy; therefore, it is clear that higher values of the pressure in the interactionzone can be achieved with higher values of the total pressure (or additionally giving anincidence to the model in the test chamber). In the same figure (right) it is reported the wall heat flux distribution for the samecomputations; in this case both the total pressure and total enthalpy influence wall heattransfer, that increases as these two variables increase (following roughly the dependencyupon the product p 0 H 0 ). 8000 1E+06 H0=11.90 P0=2.45 800000 H0=11.90 P0=2.45 6000 H0=15.00 P0=2.45 H0=15.00 P0=2.45 H0=18.80 P0=2.40 H0=18.80 P0=2.40 Heat Flux (W/m2) H0=10.40 P0=4.70 H0=10.40 P0=4.70 Pressure (Pa) H0=15.00 P0=5.20 H0=15.00 P0=5.20 H0=18.80 P0=5.20 600000 H0=18.80 P0=5.20 H0=11.00 P0=7.90 H0=11.00 P0=7.90 4000 H0=15.00 P0=7.90 H0=15.00 P0=7.90 H0=18.80 P0=7.90 H0=18.80 P0=7.90 400000 2000 200000 0 0 0 0.2 0.4 0.6 0.8 0 0.2 0.4 0.6 0.8 X (m) X (m)Figure 8. Wall pressure (left) and heat flux (right) distributions in PWT operating conditions. In the recirculation region there is a decrease of the heat flux, typical of fully laminarinteractions, followed by an increase on the flap and a peak just immediately after thereattachment point, where boundary layer thickness reaches the minimum value.
14 M. Di Clemente, E. Trifoni, A. Martucci et al. For some of the computations, the effect of the wall temperature has been estimated byconsidering the radiative equilibrium condition. The results of these computations arereported in Figure 9 for pressure (left) and heat flux (right) wall distribution. As general trend,with the radiative equilibrium temperature at the wall, the separation bubble is larger (exceptfor the condition PWT-3 where the effect is negligible) and the peak loads over the flap (boththermal and mechanical) are lower than those predicted with a fixed wall temperature equal to300K. This is due to the higher temperatures in the boundary layer in the case of the radiativeequilibrium condition, and then to the lower values of density causing an increase of theboundary layer thickness; the upstream propagation of pressure disturbances is enhanced inthe case of radiative equilibrium and, consequently, an early separation is predicted. The effects on mechanical loads is a reduction of ∼4% with the condition of equilibriumradiative wall whereas is ∼3% for thermal loads as reported also in Table 2. It can beconcluded that in these conditions surface temperature has only a small effect on thermal andmechanical loads acting on the flap. 1600 400000 H0=11.90 P0=2.45 T=300K H0=11.90 P0=2.45 T=300K 1400 H0=11.90 P0=2.45 T=Tradeq H0=11.90 P0=2.45 T=Tradeq H0=18.80 P0=2.40 T=300K H0=18.80 P0=2.40 T=300K H0=18.80 P0=2.40 T=Tradeq H0=18.80 P0=2.40 T=Tradeq 1200 H0=15.00 P0=5.20 T=300K 300000 H0=15.00 P0=5.20 T=300K H0=15.00 P0=5.20 T=Tradeq H0=15.00 P0=5.20 T=Tradeq H0=11.00 P0=7.90 T=300K Heat Flux (W/m2) H0=11.00 P0=7.90 T=300K H0=11.00 P0=7.90 T=Tradeq H0=11.00 P0=7.90 T=Tradeq Pressure (Pa) 1000 H0=18.80 P0=7.90 T=300K H0=18.80 P0=7.90 T=300K H0=18.80 P0=7.90 T=Tradeq H0=18.80 P0=7.90 T=Tradeq 800 200000 600 400 100000 200 0 0 0.2 0.4 0.6 0.8 0.2 0.4 0.6 0.8 X (m) X (m)Figure 9. Twall effects on wall pressure (left) and heat flux (right) distributions. Table 2. Tw effects: comparison of the peak values on the flap Tw=300K Tw=Trad.eq. Ppk qpk Ppk qpk (Pa) (kW/m2) (Pa) (kW/m2) PWT – 1 331.7 78.8 306.5 72.9 PWT – 3 328.8 131.6 328.7 129.7 PWT – 5 680.4 171.2 639.8 167.0 PWT – 7 1093.7 146.9 1054.3 142.8 PWT – 9 949.0 293.2 907.2 286.5
Design, Execution and Numerical Rebuilding of Shock Wave… 153.3. Definition of PWT Model The wide amount of CFD results obtained in different PWT conditions has permitted thedevelopment of the extrapolation-from-flight procedure: it allows to determine theexperimental test conditions (P0, H0 and model attitude) able to duplicate the representativemechanical and thermal loads ahead and over the flap. However, in order to give the finalrequirements for the detailed model design and then for the execution of the tests, it has beennecessary to consider also different aspects of the problem, not only the aerothermodynamicones. A detailed numerical analysis has been carried out to analyse the effects of geometricvariation of the model on the flow variables, in particular, the effects of the nose radius, theflap dimension and the model’s finite span have been considered. Sensitivity analysis hasbeen carried out considering the PWT operating condition characterized by a reservoirenthalpy H0=15MJ/kg and a reservoir pressure P0=10 bar, being this condition the onedetermined for the duplication of the point P1 flight conditions over the model as it will bedescribed hereinafter.3.3.1. Nose Radius Computations with the radiative equilibrium wall assumption have shown thattemperature in the nose region could reach 2000 K. If there will be the possibility to have anactive cooling system (at least in the nose region) the size of the nose could be decreased inorder to not exceed the model weight limit for the “Scirocco” Model Support System (MSS). A sensitivity analysis to the nose radius has been then carried out for one of the selectedoperating conditions inside the PWT operating envelope, by considering three differentmodels with the same length of the plate ahead the flap and three different nose radii, equal to0.25 m (the first hypothesis), 0.1 m and 0.05 m; it has been found that the influence of noseradius is small in terms of mechanical loads (see Figure 10, left) even if a slight decreases ofabout 5% is predicted in the reference and peak values whereas, for what concerns thethermal loads (see Figure 10, right ), a slight increase of the values in front of the flap and asmall decrease of the peak values is predicted as also reported in Table 3. 8000 H0=15 MJ/kg P0=10 bar AoA=12 deg H0=15 MJ/kg P0=10 bar AoA=12 deg 1.5E+06 6000 Rnose = 25 cm Heat Flux (W/m2) Rnose = 25 cm Rnose = 10 cm Pressure (Pa) Rnose = 10 cm Rnose = 5 cm Rnose = 5 cm 1E+06 4000 500000 2000 0 0.2 0.4 0.6 0.8 0 0.2 0.4 0.6 0.8 X (m) X (m)Figure 10. Effects of nose radius on wall pressure and heat flux.
16 M. Di Clemente, E. Trifoni, A. Martucci et al. Table 3. Nose radius effects: mechanical and thermal loads ahead and over the flap Rnose Pref Qref Ppk qpk (m) (Pa) (kW/m2) (Pa) (kW/m2) 0.25 1024.91 120.85 2226.34 266.34 0.10 1002.02 152.36 2095.86 252.26 0.05 890.92 167.18 2004.07 239.16 The reduction of the nose radius causes a decrease of the separated region mainly due tothe movement towards the flap hinge line of the separation point whereas the reattachmentpoint is located more or less in the same position for all the analyzed configurations (seeFigure 11). H0=15 MJ/kg P0=10 bar AoA=12 deg 0.03 Skin Friction Coefficient Rnose = 25 cm 0.02 Rnose = 10 cm Rnose = 5 cm 0.01 0 0 0.2 0.4 0.6 0.8 X (m)Figure 11. Nose radius effects: skin friction distribution. From this analysis it results that the model with the nose radius equal to 0.1m seems to bethe best solution for the model configuration, also considering the fact that a lower value ofthe radius could make difficult the handling and positioning of model instrumentationwhereas the model with the biggest value of the nose radius could result in a too heavy modeldifficult to sustain during the test execution with the MSS.3.3.2. Flap Dimensions Another variation that has been considered with respect to the preliminarily selectedmodel has been done by considering the full scale flap dimensions, thus exploring thepossibility to test in PWT “Scirocco” the actual EXPERT open flap before the flight, whoseoverall dimensions are 0.30 m in length and 0.40 m in width. The effect of this variation has been examined with respect to the model with the noseradius of 0.10 m, considering the same total length of the previous one since the extension ofthe flat plate has been decreased from 0.35 m to 0.20 m. The results are shown in Figure 12;considering the full scale EXPERT flap the size of the separation bubble decreases and thethermal and mechanical loads over the flap increase. The effects in terms of wall pressure areevident (Ppk increases of ∼24%) whereas are modest in terms of heat flux (qpk increases of∼5%).
Design, Execution and Numerical Rebuilding of Shock Wave… 17 H0=15 MJ/kg P0=10 bar AoA=12 deg 8000 1200 Flap 1:2 1000 6000 Flap 1:1 Heat Flux [kW/m ] 2 800 Pressure [Pa] pressure Flap 1:1 pressure Flap 1:2 600 tot. wall flux Flap 1:1 4000 tot. wall flux Flap 1:2 Geometry 400 200 2000 0 0 0.2 0.4 0.6 0.8 1 X [m]Figure 12. Effects of flap dimensions.3.4. Final Configuration and Materials The final configuration of the model, whose characteristic dimensions are Rnose = 0.10 m,Lplate = 0.20 m (the flap hinge is located at X=0.30m starting from the nose), Lflap = 0.30 m(projection on the X-axis), corresponding to the full scale 1:1 flap and flap deflection angle =20 deg is shown in Figure 13. MATERIALS 0.4 Nose : TBD Plate : PM1000 Flap : C-SiC 0.3 Lplate = 0.20 m Lflap = 0.30 m (scale 1:1) 0.2 Y (m) δflap = 20 deg 0.1 R nose = 0.10 m 0 -0.1 0 0.1 0.2 0.3 0.4 0.5 0.6 X (m)Figure 13. Final model configuration. The final geometrical configuration of the model to be tested in the plasma wind tunnel“Scirocco” is a trade-off between the aerothermodynamic requirements necessary toreproduce the flight characteristic parameters of the interaction in PWT conditions, and thethermo-mechanical design issues that have taken under consideration also different aspects ofthe problem. The model reproduces the EXPERT capsule flap (scale 1:1) characterized by 20 degdeflection angle; it is mounted on an holder with a flat plate ahead the flap with roundedleading and lateral edges. In Figure 14 it is reported a schematic representation of the model.To be consistent with the EXPERT capsule, the model will be built by using as much aspossible the same materials to manufacture its different parts: the leading edge is a GLIDCOPAL-15 copper cylinder with an active cooling system; the upper part is covered by a flat plateof PM1000 equipped with pressure taps, thermocouples and combined heat flux/pressure
18 M. Di Clemente, E. Trifoni, A. Martucci et al.sensors; the flap is covered by a 4mm thick plate of C-SiC with a deflection angle of 20 degwith respect to the flat plate, and it is equipped with pressure taps and thermocouples. Thelateral rounded panels, the entire lower panel and the parts below the PM1000 flat plate willbe realized in PROMASIL 1100; the wedge underlying the C-SiC flap will be realized inamorphous carbon. For what concerns the dimensions of the model, the cylinder leading edge has a radius of100mm and a length of 400mm, the flat plate is 400m wide and 200mm long, the flap is400mm wide and 300mm long. All the lateral edges are rounded with a radius of 50mm inorder to avoid localized over heating, whereas the flap plate has a radius of curvature at thelateral edges equal to 4mm (i.e. its thickness). The model will be installed on the PWT Model Support System (MMS) by means of aproper interface that consists of a commercial steel circular beam built with AISI 316L; theinterface is covered by proper thermal insulator of PROMASIL 1100 to avoid any criticalsolicitation due to the plasma interaction with the model surface. Such a covering has acylindrical shape for the proper alignment of the upper and lower parts of the test article withthe MSS body surface. It is realised to avoid the presence of gaps between the surfaces andthe possibility of any peak heating occurrence.Figure 14. Model for PWT tests. 4. EXTRAPOLATION FROM FLIGHT PROCEDURE The definition of representative experiments in PWT has been done by considering themost interesting points of the EXPERT reference trajectory: point P1 (M∞=13.40, h=37Km),characterized by the highest stagnation point heat flux, and point P2 (M∞=18, h=50Km)characterized by high heat flux and a relatively low pressure, potentially critical forpassive/active oxidation transition of the C-SiC.4.1. Facility Operating Conditions In par. 0, requirements for the execution of the PWT test campaign will be shown:according to the extrapolation from flight procedure, those requirements must be duplicated
Design, Execution and Numerical Rebuilding of Shock Wave… 19inside the facility on the representative model that has been defined and dimensioned. To thispurpose, it is necessary to define the facility operating conditions and model positioning andattitude within the test chamber to achieve the goal. The wide amount of CFD results obtained in different operating conditions has permittedthe development of the extrapolation-from-flight procedure, in such a way to determine theexperimental test conditions (P0, H0 and model angle of attack) that allow for the duplicationof the representative mechanical and thermal loads ahead and over the flap of the model. For each of the computations carried out in PWT conditions, considering the effect offacility operating conditions, of the angle of attack of the model, wall temperature, radius ofthe nose and length of the flap, different variables of interest have been analyzed: • Boundary layer edge variables at the separation location: M, Re, V and ρ • Reference values at the separation location: χ, V, Pref and qref The comparison between the characteristic SWBLI parameters estimated in flightconditions and the results obtained in PWT for the selected test conditions, is shown, in termsof reference pressure (Pref) and heat flux (qref), in Figure 15 and Figure 16, respectively. Forthe range of trajectory around the maximum stagnation point heat flux (Point P1, H0∼13.2MJ/kg), it seems not possible to duplicate the reference values of pressure and heat flux whichinstead could be well enough duplicated for points at higher enthalpy, that it clearly means athigher altitudes (Point P2, H0∼18 MJ/Kg, h∼50 Km). 2.50E+05 CFD Flight 2D Qref Flight 2D Qref Flight 3D Point P1 2.00E+05 Point P2 Qref - PWT Qref - PWT Additional Runs 1.50E+05 Qref (W/m2) 1.00E+05 5.00E+04 0.00E+00 0.00E+00 2.00E+06 4.00E+06 6.00E+06 8.00E+06 1.00E+07 1.20E+07 1.40E+07 1.60E+07 1.80E+07 2.00E+07 H0 (J/Kg)Figure 15. Comparison between qw ahead the flap in flight and PWT conditions. 1.00E+05 CFD Flight 2D Pref - FLIGHT 2D Pref FLIGHT 3D Point P1 1.00E+04 Point P2 Pref - PW T Pref - PW T Additional Runs Pref (Pa) 1.00E+03 1.00E+02 1.00E+01 8.00E+06 1.00E+07 1.20E+07 1.40E+07 1.60E+07 1.80E+07 2.00E+07 H0 (J/Kg)Figure 16. Comparison between Pw ahead the flap in flight and PWT conditions.