Your SlideShare is downloading. ×
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
mario felipe campuzano ochoa / parametric std. @ ramp flow
Upcoming SlideShare
Loading in...5
×

Thanks for flagging this SlideShare!

Oops! An error has occurred.

×
Saving this for later? Get the SlideShare app to save on your phone or tablet. Read anywhere, anytime – even offline.
Text the download link to your phone
Standard text messaging rates apply

mario felipe campuzano ochoa / parametric std. @ ramp flow

2,023

Published on

Published in: Education, Technology
0 Comments
0 Likes
Statistics
Notes
  • Be the first to comment

  • Be the first to like this

No Downloads
Views
Total Views
2,023
On Slideshare
0
From Embeds
0
Number of Embeds
1
Actions
Shares
0
Downloads
1
Comments
0
Likes
0
Embeds 0
No embeds

Report content
Flagged as inappropriate Flag as inappropriate
Flag as inappropriate

Select your reason for flagging this presentation as inappropriate.

Cancel
No notes for slide

Transcript

  • 1. National Aeronautics andSpace AdminishationAdvanced Research X-Program DivisionCenter Confidential and Secret Level Category X-33Washington, DC 20231USA Mixing and Combustion in Hypersonic Airbreathing Propulsion Systems Mario Felipe Campuzano Ochoa NASA FellowNGT-70380 Hypersonic Airbreathing Propulsion Branch Gas Dynamics Division NASA Langley Research Center August, 1996
  • 2. 7 Contents 1 Introduction 3 2 The Airframe-Integrated Scramjet . 3 2.1 High-Speed Propulsion Characteristics 2.2 MixingEnhancement- Streamwise-Vorticity b Modelling of Reacting Internal Flowfields for a Scramjet Erigine I 3.1 MathematicalModeling 8 4 Flow Losses and Thrust Potential in Scramjet Combustors T4 5 Numerical Simulation of High-Speed Reacting Flows - Hz - Air 15 rt Concluding Remarks 18
  • 3. 7 Abstract A brief presentation of the basic characteristics of hypersonic airbreathing piopulsion systems will be conducted. The focus of the research project wili be on supersonic/hypersonic reacting internal flowfields. Including a detailed chemicai kinetic mechanism for the combustion of hydrogen and air, and the performance analysis of a SCRAMjet engine. Also the numerical simulation of a high-speed reacting flow wiil be pre- sented with a brief analysis.
  • 4. Symbols eA; constant in Arrhenius law . bt, body force of species i in x coordinate direction . b;a body force of species i in y coordinate direction oCt Concentration of species i , D;i binary diffusion coefficient oDr thermal diffusion coefficient oE; activation energy; total int. energy .f; mass fraction of species i oGn Gibbs energy of reaction . 9;, gibbs energy of species i oh; enthalpy of species i ,h? reference enthalpy at standard conditions oK equilibrium constant ok thermal conductivity oku reverse reaction rate .kt forward reaction rate tM molecular weight ol[ constant in Arrhenius law fol reaction i a Ttt moles of species i .7, effective temperature oilt streamwise diffusion velocity of species i
  • 5. diffusion velocity of species itransverse diffusion velocity of species i mole fraction of species iratio of specific heats, stoichiometric coefficientequivalence ratio
  • 6. IntroductionBfficient air-fuel mixing rate is critical for the successful operation of hy-pervelocity and scramjet engines. Even if the best combination of mixingenhancement mechanisms is used a considerable loss in engine thrust is ex-pected. These particular engines are extremely sensitive to all type of lossespresent in this complex environment[3], which includes, complex shock andexpansion waves) three dimensional flows, complex chemical kinetics, abruptheat tranfer rates, non-equilibrium phenomena, and many others sub-aspectsof the above examples. Enhancing the mixing of fuel and air, wouid allow, for the diffusiue-scramjet, shorter combustors with consequent reduction in weight, heat trans-fer, and shear losses. However, the key of efficient mixing is to enhancemicromixing (small scale mixing) via techniques which do not increase dra-matically the overall combustor losses. Furthermore, to achieve micromixingis a difficult problem since the airflow moves very fast through the engineand there is very little time for fuel to mix with the air and burn efficiently.For example,, at high-speeds such as Mach 7, a fuel particle stays in thecombustor for less than a millisecond.2 The Airframe-Integrated ScramjetThe NASP (National Aerospace Plane) concept relies heavily on the integra-tion of the engine with the entire aerodynamic geometry[2] of the vehicle,specially with the undersurface of the plane. This surface must be designedwith caution, so that the engine, the forebody undersurface, and the after-body, all function as a group, meaning each part is heavily dependent on theother. The NASP concept is defined as a horizontal takeoff and landing (HTOL)type of vehicle. Were the undersurface of the vehicle is involved in compress-ing and expanding the incoming flow-stream. The reason this type of designis used is mainly because to achieve hypersonic velocities, a considerableamount of thrust is required, and the typical physical size of a scramjet willnot capture all the airstream needed by itself. Therefore the forebody of theplane wili serve this function and act as part of the inlet of the scramjet. The fuel of choice is hydrogen, which burns very rapidly compared with
  • 7. the hydrocarbon fuels used in conventional jets. The hydrogen would bestored in liquid form to minimize fuel tank size, but would be injected ingaseous form to speed combustion. Note that although the hydrogen fuelis characterized by its high-energy potential, it is not enough to supply therequired energy for the needed thrust, therefore large amounts of free-streamair are critical for the successful operation of a scramjet engine. In high-speed flow a bow-shock is produced by the vehicle undersurfacesection (figure 1). This shock wave compresses the incoming air and helpsto reduce the compression required by the engine inlet, therefore a smallerengine can be efficiently used and the forebody undersurface is responsiblefor helping capture the necessary flow-mass. To take advantage of all theair compressed by the forebody undersurface, a series of engine moduies areplaced along the spanwise undersurface to ingest as much compressed airas possible. One negative aspect of this design is that the thick boundarylayer along the undersurface of the vehicle is ingested by the scramjet causingshear and viscous losses. The aft portion of the aircraft underbody would beshaped to serve as the upper half of the engines nozzle. A dual combustion mode scramiet is needed to accelerate the vehicle fromtake-off speed to hypervelocity speeds. For example one possible arrangementwill be to acomodate a turbojet engine on top of the scramjet (figure 2).This turbojet will be responsible for the takeoff-Mach3-4 propulsion regime.Above Mach 4, the turbojet engine is turned off via a mechanical device (i.e.adjustable door) and the scramjet engine ignites and becomes responsible forthe propulsion of the high mach number regime. A door to close the scramjetbelow Mach 4, can also be utiiized but it may not be required. Note thatthis is just an example of many dual-mode combustion systems that havebeen proposed. The most recent version can also be seen from the bottomof figure 2[16]. A common scramjet engine will look like the figure 3. As can be seenthe engine components[7], are a cowl that helps to pre-compress the flow,several fuel-injection struts which provide locations for tangential and normalinjection of gaseous hydrogen, sidewalls, a very short combustor, and a nozzle.The sidewalls are sweep to an angle, so that stream flow can be capturedefiiciently by the engine, since the flow is turned by the sweep angle of thesidewalls a logic arrangement will be to position the fuel-injection struts alongthe same angle, so that losses are minimized (i.e shock losses generated atthe leading edge of the fuel injection strut can be significant compared to the
  • 8. overall engine losses). The cold hydrogen is also utilized as a coolant, Regenerative coolingcombined with modern high-temperature structural technology makes thestructural design of such an engine feasible if the design minimizes movingparts exposed to high-temperature flow. Once the hydrogen absorbs all theheat, it is iniected. This cycle optimizes the process and the conservation ofenergy principle works in favor of the efficiency of the engine. In general the design of hypersonic airbreathing propulsion systems ishighiy integrated with the airframe. And the combustor is responsible for aconsiderable fraction of the friction, heat transfer, and mixing losses.2.L High-Speed Propulsion CharacteristicsT.ypical operating conditions for a scramjet engine are: r Velocity(mean): The order of free stream r Pressure: one-half atmosphere o Temperature: 2,000 to 4,000 Rankine (1000-2500 Kelvin) o Mach Number: Il3 of free stream The fuel injection strategies employed, are typically normal and/or tan-gential to the freestream, At the low end of the mach number (6-12) thehydrogen can be injected normal and parallel to the free stream. However arecognized problem for the mid-range mach number (6-12) is the relativelylow static temperatures present in the combustor (i.e this temperatures candelay ignition). At higher mach numbers, thrust generation becomes a chal-lenge and normal fuel injection can not be utilized, since all sources of lossesmust be minimized, ideally to zero. Normal fuel injection causes a bowshock (figure 4) in front of the jet and consequentiy increases mixing andlosses substantially . However at the high Mach number end this losses cannot be tolerated, and fuel must be injected in a parallel fashion. Figure 5shows the physical locations of the fuel injection with respect to the engine. Since the hydrogen is used as a coolant its temperature increases sub-stantially and the injection velocity can reach values near 12,000 ft./sec. orgreater, having a dramatic effect on the mixing rate[9]. This means that the
  • 9. mean shear between the fuel jet and local stream flow becomes small in theMach I2-lS lange due to the high effiux speed of the regeneratively heatedhydrogen fuel, further reducing the innate shear-induced mixing. Thereforein terms of criticality and payoff, the high Mach number range is the mostimportant(air-breathing) regime for mixing enhancement.2.2 Mixing Enhancement - Streamwise-VorticityStreamwise vorticity is a powerful technique and currently the preferred can-didate to mix efficiently fuel and air in a scramjet combustor. One form ofgenerating this type of vorticity is by the use of fuel-ramp injectors. Thesefuel ramp injectors are primarily used to alleviate the problem of slow mixingof fuel and air when parallel injection is used. As discussed previously paral-lel injection is required at high Mach numbers to generate a net thrust andalso to extract energy from the heated hydrogen. The fuel-ramp devices willinduce axiai vorticity and local recirculation regions. This recircuiation re-gions can represent a rearward-facing step commonly used for flame holdingin high-speed reacting flows. Streamwise vorticity can be used as an efficient mechanism for mixing,however macro-mixing produced by the vortices should lead to efective micro-mixing (small-scale turbulence) to allow a proper level of combustion. Alsoshock-vortex interactions can increase mixing by a factor of three, whilestreamwise vortices can increase maclo-mixing by a factor of two. Schock-mixing layer interactions can have an influence up to forty percent. Com-bustion usually does not affect mixing substantially, however, turbulence can dramatically increase chemical reaction[6]. In general mixing enhancement for engine performance improvement is the degree of increased micromixing(small-scale turbulence/molecular mix- ing), increased contact area between air and fuel. The following physical parameters have a significant impact/effect upon turbulent shear layers, and therefore are strong candidates for mixing enhancement tools[10]: o Pressure gradients o Flow curvatures o Energy release
  • 10. Shock/expansion wavesInjection locationsIt must be added that the flow losses associated by placing a fuel-rampinjector in a supersonic combustor are not disregarded. In fact a care-fully balance analysis should be made between the gain in combustionefficiency due to the mixing incurred by the streamwise vorticity andthe loss in thrust due to the cross-stream flow introduced in the com-bustor. The flow losses for a fuel-ramp injector include: - Additional Shocks - Frictional Drag - Pressure drag on the face of the ramp - Vorticity generation losses - Base pressure - Flow recirculation (rearward-facing step)3 Modelling of Reacting Internal Flow-fields for a Scramjet EngineSupersonic/Hypersonic internal flow fields are extremely complex. Inthese flow fields there is a major synergism (interaction, such that thetotal effect is greater than the sum of the individual effects) betweenthe fluid dynamics(inviscid and viscous), and chemical reactions takingplace.The exact effects of chemical reaction processes are not well understood.Chemical reactions produce species in the flow that affect transportprocesses, energy can be released or consumed by chemical reactions,and turbulence can affect concentration of species that can alter steady-state reaction mechanisms[5]. Therefore a key aspect of hypersonicpropulsion is the understanding of the interaction of viscous turbulentflow fields with exothermic (i.e. formed with evolution of heat) chemicalreactions and transport processes.
  • 11. 3.1 Mathematical ModelingChemical reactions present in a scramjet combustor can be studiedand analyzed by modeling high-speed reacting mixing iayersf11]. Anexplanation for this is that the combustor flowfield can be viewed as acollection of spatially developing and reacting high-speed mixing layersor iets from fuel injectors mixing with air.The two-dimensional Navier-stokes equations can be coupled with thespecies continuity equations. The following equations are for a two-dimensional case (2 D) Navier-Stokes, energy, and species continuityequations governing muitiple species fluid undergoing chemicai reaction[1].The finite rate chemical reaction of gaseous hydrogen and air can bemodeled with a seven-species, seven-reaction model, or a more com-plete mechanism of nine-species, eighteen-reaction model. The coeffi-cients governing the momentum, energy and mass are determined frommodels based on kinetic theory. Sutherlands law is employed to com-pute the individual species viscosity; and when a mixture is presentWilkes law is used. The species thermal conductivity is computed us-ing a modified version of Sutherlands iaw. And the mixture thermalconductivity is computed using Wassilewas formula. For the trans-port properties, the Chapman and Cowling law is used to determinethe binary-diffusion coefficients which describe the diffusion of eachspecies into the remainning species. Knowing the diffusion coefficients,the diffusion velocities of each species are determined by solving themulticomponent diffusion formula.After the thermodynamic properties, chemical production rates, anddiffusion coefficients have been computed, the governing equations aresolved using a general conservative numerical approach. dU*aF(u) *Eg(u) :u (1) 0t 0r dywere the vectors are:
  • 12. p p pu - o, PU2 U- pu ) F_ puU - Tyr pE (pE - o,)u r,yu * Q, pf; puf; * ptr;f; pu 0 ouD - T^^, a9 PD f ;b; ":l I -pu2 oy H- PL f;b;o (pE-or)u-rr"uIqE pD ftb;(Y -p Vi) t pu fi -t p6;f ; ti;;the terms i nside the vectors are: or:-p*l( 0u 0u. 0u or f oy)i-2tt^ (2) ^ ^ or oa:_p+^(#*H,*rr# (3) Tro : Tsr: ,(X * Hl (4) n, : -k#*,i h;f;t t + Rrntffi)@,. - il,i) (5) ry ry" AT pL,hofpo R"TLo_L,=r(ffi)(tr . * ,X;Dr;,, - Qa: -r u, * + - ,r) (6) E:S h;f;-P-*u2+z 2-02 (7) z=l rT 4:h?* JTa cp;d.T I (8)
  • 13. N"t p: pR"T#, (e)The diffusion velocities can be found from: YXt- Y,Wvi-v)+(ro-uY* j=1 li r,r,ot - bi, . *r r#,t+ - ff x$ t (10)Given a system with a number of chemical species, lys, then f :1,2,..., (l/r - 1) and -l/s - 1 equations must be solved for the species fi.The resulting mass fraction can be found from the mass conservationlaw: Lf;:t i=l (11)Solution Procedure - Thermodymamics 1. The first step in solving the above system of equations is started by finding the specific heat of every species present in the chemical reaction, which can be found from a simple fourth order polyno- mial in temperature T. + : R A.; * B;T * C$2 I D;73 + E,Tn (r2) The coefficients can be found by a curve fit of data available. 2. Then the enthalpy of each species is found from equation 8. , J. Finally, the total energy of the system can be found from equation 7. 10
  • 14. 4. Each chemical reaction present in the reacting system is charac- terized by an equilibrium constant, which can be found from the Gibbs energy of each species. Given a constant pressure process the Gibbs energy of each species is: D;r^ 2 :,q,ff -Ttnr)- *r R -1 ----/ 2- -?r" - 6 12 - *r 20 *4 -GT (13) The last two coefficients can be found from a reference giving the thermodynamic properties of the corresponding element and the Gibbs energy of reaction can be found from, the reiation: LGp: t 2= pr oduc ts n;Lg; - I eact n;Lgt (14)5. The next step will be to find the equilibrium constant for each reaction: 1 " , -ACn, /r:(p.,-7)o"eree*f) (15) and An represents the change in number of moles from reactants to products. Solution Procedure - Chemical Kinetics Chemical kinetic effects can have a dramatic effect in high-speed combustion systems. Because of the conditions that exist in high- speed combustors, particularly short residence times, chemical ki- netic effects can be very important. In an ideal combustor the combustion process would be mixing controiled. For the combus- tion process to be mixing controlled, the reaction rates in the com- bustor must be very fast, however in a high-speed combustor the short residence times and the low pressure can create combustion problems which prevent the achievement of the ideal situation. Chemistry makes itself felt in reacting turbulent flows through the chemical source terms that appear in the equations for species and energy conservation. The main goal here is to obtain a quantity 11
  • 15. tl;, defined as the mass rate of production per unit volume ofspecies i.Considering a system involving N chemical species, and 1y6 re-action steps. The forward rate of each reaction j, is given by theArrhenius iaw: kri: A"r|erp(-#) (16)The next step will be to find the backward rate: L.. koi: ?/ T] (t7)The subscripts / and b on k identify rates for the reaction occur-ring in the forward and backward directions, i.e., in the directionsindicated by the arrows in the following equation. K is the equi-librium constant from above.Knowing the above quantities, the production rates can be foundfrom the law of mass action: N" N,. DtinCo + t ti:,;Cr, (i : r,2, ..,Nn) (1s) i=l i=1And the arrows denote the rate constants that depend on lemper-ature for reactions in ideal gas mixtures.The rate of change of concentration of species i by reaction j isgiven by: (r: - v z )s - 0,lo - ti)@rifr,ri" -,ro,flci,i,) (te) i=-J, i=lobserve that the sign of the gammas automatically accounts for thecreation of products and destruction of reactant" (^li!,n representsthe product, and 11; the reactants).The rate of change in concentration of species i by reaction j isfound by summing the contributions from each reaction 1,2
  • 16. -A/p ci : L(ci) j (20 ) j=1The production of each species i can be found from: dst: CtMi (2r)Solution-Procedure Molecular Diffusion ModelsThe individual species viscosities are computed from Sutherlandslaw, u | ,T.s.T,+S I t. (22) po-ToT+s -where po and To ane reference values, and S is the Sutherlandsconstant. As soon as the viscosity of each species is calculated themixture viscosity can be computed from Wilkes law: _A,/s pi F^:D r -f u1 i:l | rlXnLf:, Xidn, (23)where: (1 + ((ffx lDt(ffi)+Y .t v;j - (24) fto + ffit+The species thermal conductivities can be calculated from Suther-lands law (but with different values for the reference values): k ,T .e"7"+ SI l. (25) kr, ToT+S -and the mixture thermal conductivity is computed from Wassil-jewas formula: l/" ki k^:D I+IlXiLl:rXidoi (26) i=7 13
  • 17. where doi : I.065$i1. The binary diffusion coefficient between species i and j can be found from the Chapman and Cowling kinetic theory result: n,ij 0.00185873 (#,^+ fi r,t (27) "roipro The diffusion collision integral is approximated by: c*.r6r LDD :7*ota5 + (f. + 0.5)-2 (28) where T* : rL. NlOnce ali the binary difusion coefficients for ali the species present are know,the diffusion velocities of each species can be found. The diffusion velocity isthe species velocity due to all diffusion processes algebraically added to theconvection velocity.4 Flow Losses and Thrust Potential in Scram- j"t CombustorsA logic way to measure performance mixing efficiency is to compare a scram-jet engine with no mixing enhancement, with a forced mixed combustor. Ifthe thrust of the forced mixed combustor is greater then the design is accept-abie. However if the thrust obtained is smaller then the design is obviuoslynot acceptable. However, it has been verified that an excellent and betterparameter for measurement of flow losses is the non-dimensionalized com-bustor effectiveness parameterf4], which is basically a measure of the thrustpotential of a given flow-field with a particular inflow or the useful potentialwork measured from some reference state. Where for a Scramjet engine thereference state is the atmospheric pressure, and the potential work is simplythe thrust power potential of the flow when expanded isentropically to thereference state.To start an efficient analysis, the thrust can be found from: r: | (ou + P)dA" - | Q" + P)dAi (2e) 14
  • 18. where i is the inlet and e is the exit. For an ideal nozzle performance, thepressure at the exit is taken as atmospheric. From this definition an efiiciencyparameter can be written as the ratio of the power terms(for a given flightvelocity U). (T IJ )o,tuot Tactuat ttce ( 30) (T Ll ) ;a""t T;,t,ot by isentropically slowing the combustor inflow to aT;s"o1 ca.rr be obtainedMach number of zero, allowing the fuel to react to chemical equilibrium atconstant pressure) then isentropically expanding the flow to the ambient pres-sure and computingthe stream thrust at the nozzle exit. To obtain the idealthrust the inflow stream thrust is substracted from lhe nozzle stream thrust.Tactuat is calculated in a similar way, but now the nozzle exit stream thrustis obtained by an ideal expansion from any combustor station to ambientpressure.5 Nurnerical Sirnulation of High-Speed Re- acting Flows - H2 - AirA computational study of the effect of fuel and air temperature and fuel andair velocity is performed in this section. It is important to know the effectsof gas temperature on supersonic combustion since the hydrogen will be usedas a coolant and therefore its temperature will increase substantially beforeinjection takes place. Most previous studies have involved room-temperaturefuel injection, various authors (ref 12 and 13) have not found any physicaltrends on the main effect of fuel temperature on mixing rate.The numerical simulations performed, simulate the case of two freestreamflows (air to hydrogen ), where the air is composed of 23.14 percent of oxigen,and 76.86 percent of nitrogen. A case was selected to serve as a baselinemodel, and two other cases(case 1 and case 2) were compared to this baseline model, to see how the change of reference conditions wiii alter thereaction production of water.For all cases the ratio of specific heats was taken constant ) : 1.4, thefinite rate chemistry model was selected, and the no. of chemical specieswas seven with seven chemical reactions. The laminar Prandtl and Schmidt 1r Ir)
  • 19. numbers were 0.72 ald 0.22 respectively. The turbulent Prandtl and Schmidtnumbers were 0.9 and 0.9. The two-equation non-algebraic turbulence modelof Menter was selected. The mesh size for the jets is 65x130 and for thereacting domain downstream is 130x65, with grid clustering along the plate,and assuming turbulent flow along the plate.The boundary conditions were held fixed at the inlet at the reference valuespecified, and an adiabatic wall condition were imposed on the flat plate thatseparates the two flow streams and starts at x : 0 and ends at x : 0.5 . Atthe top and bottom of the domains an extrapolation boundary condition wasimposed and a zeroth order extrapolation at the outlet was aiso imposed atthe outlet.The figures are defined as follows: - case 1 (A and B) where B is the baseline - case 2 (C and B) where B is the baseline - for xy figures case A, B(baseline), CCase 1-B baselineThe reference air conditions for the baseline case 1-B are an incoming Machnumber of 1.9, with a temperature of 1000K, and a density of 0.35151(glmt,with an air velocitv of 1200 mf s. And for the hydrogen a fuei velocity of3800rn/s, with a temperature of 1000K, and a density of 0.02456Kg1-t.From the temperature contour figure, it can be seen that the temperatureremains uniform along the flow fieid except at the region near the middle ofthe computationai domain were most of the reaction is taking place. Thetemperature increases dramatically near this region due to the strong chem-istry events taking place. The approximate temperature is about 2200K forthe reacting part of the flow field with a small portion reaching a value of27A0K. It can also be seen that the temperature is higher were the lowestamounts of concentrations of water are present, suggesting a large heat trans-fer rate between the gas mixture and the water. However from the xy graph(case B - Y location is at the middle of the plane) the change in temperatureand water lerrel is not dramatic, and as the temperature decreases the waterlevels tend to increase at the same rate. The ratio of the velocities is: U 1u"t :3.161 (31) Uoi, l6
  • 20. Case 1-AFor case 1-A the reference air conditions are an incoming reference velocity of2200 mf s, with a temperature of 1000K, and a density of 0.35151{ g f m3. Andfor the hydrogen a fuel velocity of. 1,800m1s, with a temperature of 1000K,and a density of 0.024561{ g l-t .From the temperature contour figure, the temperature remains uniform alongthe flow field except at the region near the middle of the computationaldomain were most of the reaction is taking place. The temperature changesare not as dramatic as in the baseline case, reaching a maximum value ofaround 27A0K. The water levels are very small compared to the baseiine case.Suggesting that the velocity difference in the two-flow fields dramaticallysuppress combustion rates. For the xy graph (case A) the pattern is thesame as for the baseline case, but obviously the production rate of water islower. The ratio of the velocities is: (32 ) W:0.818It can be said that by changing the above ratio to a number different thanunity the combustion efficiency can be improved. But the ratio should becareffuly selected since the convective Mach number (a well know parameterin the study of mixing layers) is sensitive to this type of velocity changes,and a large convective Mach number introduce compressibility effects intothe flow field that can dramatically supress combustion processes.Case 2-CFor this case the reference air conditions are an incoming reference velocityof 4200 mf s,with a temperature of 1400K, and a density of 0.3515. And forthe hydrogen a fuei velocity of 1800m/s, with a temperature of 1000K, anda density of 0.02456.As can be seen from the temperature contour figure, the temperatute re-mains uniform along the flou field except at the region near the middle ofthe computational domain were most of the reaction is taking place. Theimprovement in combustion can be noted inmediately over the baseline case.Although the difference in velocities is higher for the baseline case, the tem-perature of one of the gases was increased by 40 percent and a considerabie t7
  • 21. improvement in combustion efficiency was obtained. Note that as expectedthe amount of water present at the mid-y location is greater for this case ascompared to the baseline case. The ratio of the velocities is; U yu"t :0.427 (33 ) Uou6 Concluding RemarksThe benefits of achieving a high-performance propulsion system using anairbreathing cycle will have an impact in commercial technology and gor,-ernment applications, meaning that high engine performance will permit asubstantial reduction in propellant and tank weight, total system weight,over-all dimensions, and consequently reduction of total cost; in comparisonwith the conventional rocket systems used to reach hypersonic speeds.Much work needs to be done, specially in the area of high-speed mixing athigh Mach numbers, which currently is the most challenging problem forsuccesful operation of scramjet engines. It can be easily concluded thatCFD is essential to the succesfui analysis of high-speed reacting flows, whereexperimental facilities fail to simulate real flow conditions.The numerical simulations shows the importance of gas temperature in propermixing. In a mixing layer simulation, increasing the fuel temperature can de-crease the mixing rate by indirectly increasing the fuel velocity resulting ina reduced velocity gradient through the shear layer. 18
  • 22. -Bibliography [1] Carpenter N4.H., "A Generalized Chemistry Version of Spark", IASA CR-4196, 1988. [2] Henry J.R., Anderson G.Y., "Design Considerations for the Airframe-Integrated Scramjet", NASA TM X-2895, 1972. [3] Krishnamurthy R., Rogers R., Tiwari S., "A Numerical Study of Hypervelocity Flows Through a Scramjet Combustor", AIAA Paper 94-0773. [4] Riggins D., McClinton C.R., "Thrust Modeling For Hyper- sonic Engines" AIAA Paper 95-6081. [5] Jachimowski C.J., "An Analytical Study of the Hydrogen Re- action Mechanism With Application to Scramjet Combus- tion", NASA TP 279I, 1988. [6] Drummond J.P., Rogers R.C., Hussaini M.Y., "A Detailed f{umerical Model of a Supersonic Reacting Mixing Layer" AIAA-Paper 86-1427. [7] Holland S. D., "Internal Aerodynamics of a Generic Three- Dimensional Scramjet Inlet at Mach 10". NASA TP 3476, January 1995. [8] Voland R., Rock K., "NIASP Concept Demonstration En- gine and Subscale Parametric Engine Tests", AIAA Paper 95-6055. [9] Wendt M., Stalker R., Jacobs P., "Effect of Fuel Tenrperature on Supersonic Combustion", AIAA Paper 95-6029. [10] Kailasanath C.L., Book D.L., "Mixing Enhancement due to Pressure and Density Gradients Generated by Expansion Waves in Supersonic Flows", AIAA Paper 91-0374. 19
  • 23. [11] Mukunda H.S., Sekar 8., Carpenter M.H., Drummond J.P., "Direct Numerical Simulation of High Speed Mixing Layets", 1ASA TP_3186.[12] Lepicovski J., "Total Temperature Effects on Centreline Mach Number Characteristics of Free Jets", AIAA Journal, 28!):a78-a82, 1989.[13] Hyde C.R., Smith B.R., Schetz J.A., Walker D.A., "Turbu- lence Measurements For Heated Gas Slot Injection in Super- sonic Flow", AIAA Journal, 28(9):1605-1614, i990.[14] Heiser W.H., Pratt A.T., Hypersonic Airbreathing Propul- sion, AIAA Education Series, 1994.[15] Anderson J.D., Hypersonics and High Temperature Gas Dy- namics, McGraw-Hill, 1989.[16] Emami S., Trexler C.A., Auslender A.H., Weidner J.P.,, "Ex- perimental Investigation of Inlet-Combustor Isolators for a Dual-Mode Scramjet at a Mach Number of 4", NASA TP 3502, May 1995. 20
  • 24. r-4 0c t E (D a l" - ; i ll-, lr u^: ---r- 1 I U) o C) !...> il- v li lr 1 o 5 x ",ll-11 P d g6 rrl lr F l- =7 lr g / frc> F.. () rrrl ^ ---l- 6- 8= tt1P. = J i+ L/1 . (D og (r( r 3 :+ tlo ( )-.t.t P. oJ -J FO z. rrl z. 6) z- rf! O = { (f c g rrl -{ L/1 - F I c a -l
  • 25. ln v? .cf + 1,/" w H I O O n n @ t- rT.t I c) rrt O c) rrt = o L,/1 -{ O =_ 7t z. --t = rrrt r:rl .:r5 -{ a) -{ C q @ 6 O o FI { f o J= (o o_ o o (D
  • 26. €Ln s= G) Fi @ c) o rnT .tl ."2. 5. C) Oo r .o rn =# =) VI c= nn -rl C) rrl T a5 vl - I oc)r- Z. Or rrr .o tlt ; z. z. s -r-lri -{J c rrlli--i-i t-liti^ -4e z.il:ril n L- rrtil c) -{il c) o z.llLn.l o ;lt v1 = @ = -{ c n w -l O n c T I C -{ LA n -{ rn = rrl z. U O Tl z. o N N T- rrt
  • 27. c z (] (D a o c o T gl -m zf, J a .t o. E 9 t ci q q o 3 Dt -rl.1, r: o2 rZ --H-li r-l;li -rti HU j Ac- ll* e+ o om 3 .a 0 74 0t o o s. 62, :c- o = o :l Q o C- #fl rn o ,+ ni cr o 9o c a o. gl n6 a 2# r r
  • 28. o o €o a ) f od oo i( D(/) <T (D6 ao 5 q8 =r{ o DC o= =r. o- otD 7.?tiit 6 oii-l_c g) o IC ) CNii lr o oIr oll/ 9.rrll,II - l a JOt o 0J :6 o(t o ao a.Q t o< o = ! o o It, 3 o o f, c @ !i o l$ - Tt A c o c_ ) 5 o o o a z o o il N f o N a o. o 9) lA (o o o a & o x E 0) J (t o L .o- o il o f,
  • 29. vz FD z o o c0 3 o- L) oa t* I t o o I :- o a< tJ o N) q p (lF _o 7 a o o o- FD c0 + o t I a ,t^ o /__--/ :- C0 V o o a o- o (D c) (D CD q o {1. gio po J- * t) OO OO { :.F ;tttf!6
  • 30. Unswept ramp Swept rampFigure 11. swept and unswept ramp fuel-injector configurations.
  • 31. Lo @ N (o o cD l.- Lo c{ o (o c! o (o 6 tqq:a?q0qotr!Q9oN(o <o$N-o)l-LoIa?qqqqqO cD o Ir.- @ LO <. N r O t N @ F- LO Cr) F @ LO st Cr) N r O ol Co l* LO t c) N - FFFr@l:(oLOtCDN- o (s =o=lJ-CDc(JGoE,.= E c I .9rgCts cl (U o oEO (6o -xcoooL+r(uLoCLEoF uorlecol-A
  • 32. v o N-LO@N@$l---t/)o)ClN bOCDF-F(c)OrOO)(f)@N@-@ o d c.j d cd d c.j o N d c.,i o N rj q f Ol-+Ol-*rl-tF@<f,-NO P F- LO $ O - O o) lt- (o LO cD N - @ LO (s NNNNNNFFT-O)CO o o- E o Fo=trct)co(EoE Ea c I .9IE c(E 6lF (s (J oIoJo tr x+.cooEJ+i(uLoCLtroF uo!lecol-A
  • 33. (o@ NtOO)NIr)COFtCOF$NCt(o O-N<f,tO(o@o)ONcDt(ol-CO 6tO(oN@$OF-c9o)rtrN(D rooO@Ocr)lt-Fst@-tOo)N(o ci aqAqqqqq-oqtqa? z LOLO$++Cf)Ci)NNNFFTOO -c o (6 =-9lrct)co(Eotr EI-at* c ol .9IE m (5 (JCts oL()o tr xcooE(ELoCLEoF uorleco-l-A
  • 34. v N-LOCON@sfl-FIf,o)NN cO crj N r (O O LO O CD CO C! @ - o dcjocorj cjoN,rjAi cjNLrj @ q OF.+Ol-$FN+-@sl--NO F Lo < (9 - o o) l- @ !o Co N F @ N(jNoINo|jFTFFFF-Or@ Lo (6 o E oo=trct)co(EoE,L. EE c t* ol .o (sJ- 91 v, otrts or-oJo tr xcooE)+J(ELoCLtroF uo!lecol-A

×