A Variable Cycle Engine (VCE) can be defined as onethat operates with two or more thermodynamic cycles. It is a type of aeroengine whose thermodynamic cycle can be adjusted by changing some components’shape, size or position, and the cycle parameters, such as pressure ratio, massflow, bypass ratio and thrust. It can be varied between those of a turbojet anda turbofan, making it to combine the advantages of both. These measures mayenable the engine to obtain the optimal thermodynamic cycle, and to acquire thegood adaptability to various flight envelopes 1.

             The engine can work as the turbojet when theaircraft requires high specific thrust, such as take-off, acceleration andsupersonic cruise. It also can work as the turbofan when the aircraft requireslow fuel consumption, such as standbyand subsonic cruise. The most important advantage expected from using VCE infuture supersonic transport is a substantial range improvements as compared toa conventional engine. These range improvements are mainly achieved by reducingthe subsonic specific fuel consumption by around 15% (relative to a Turbojet)and improving the fuel consumption at off-design by the extensive use ofvariable geometry. The future VCE will have a low emission combustor andafterburner.

The noise level at take-off will be met by FAR part 36 requirement.In other words, the future VCE will be environmentally accepted 2.The disadvantages are mainly an increase in theengine weight and a more              complexcontrol system, therefore the reliability of the engine will be affected. Theperformance of any VCE depends critically on the attainment of the predictedtechnology level improvements. The purpose of research on VCE is to improveoff-design performances, in order to satisfy the needs of broad flightenvelope, large combat radius and long cruise duration 3.Thework mode of VCE discussed in this article is presented in Figure 2:Single bypass mode: The selector valveis closed and all air goes through the Core Drive Fan Stage (CDFS). The fanbypass flow bypasses the core engine through the inner bypass duct and remixeswith the core flow downstream of the low pressure turbine. The nozzle is fullopen to shift the loading to the HP shaft to cope with the added work of theCDFS.

At the same time, the expansion ratio and the flow rate rise to increasethe specific thrust with low bypass ratio under supersonic and accelerationcondition.          Double bypass mode: The selector valveis full open and the nozzle is now closed to unload the HP turbine and load theLP turbine. The bypass ratio increases for best specific fuel consumption forsubsonic cruise and best exhaust velocity conditions for improved noisesuppression on take-off 4.Axial flow compressor is one of the most importantparts of Gas turbine engine. Axial-flow compressors are used in medium to largethrust gas turbine and jet engines. The compressor rotates at very high speeds,adding energy to the airflow while at the same time compressing it into asmaller space. The design of axial flow compressors is a great challenge, bothaerodynamically and mechanically.

The aerodynamic compressor design process basicallyconsists of mean line prediction calculation, through flow calculation, andblading procedures. The mean line prediction is the first step withincompressor design. It is a simple one dimensional calculation of flowparameters along the mid height line of the compressor where global parametersas the annulus geometry, the number of stages, and the stage pressure ratiosare scaled 5. It is necessary to design axial flow compressor at preliminarylevel and require parameters can be checked at initial level so furtherimprovement can be made at primary level before start a Detailed design.It is a challenging job to design appropriatecompressor to meet the demands of VCE, which is the compressor should implementperformance adjustment of the engine and ensure the efficiency being maintainedwithin a higher range. As described by similarity principle, multipleconditions can be converted to the same working condition according to the ruleof equality in reduced wheel speed, reduced mass flow and Mach number alongcircumferential direction and so on. Then, the compressor performances underdifferent working conditions are approximately equal to each other.

In thisresearch, a high loaded high-pressure compressor with high compression ratioand large enthalpy rise was designed for VCE in two operating modes of lowbypass ratio (single bypass mode) and high bypass ratio (double bypass mode)according to this principle 3.Thecomplexities of a supersonic flow in compressor have been summarized: wavestructures such as expansive waves and compressive waves (even the shocks)exist in supersonic regions. Due to this, flow parameter changes drastically inthe channel. Compressive waves may be formed by disturbed flow in that regions,so influence on aerodynamic performances caused by compressive wave-boundarylayer interactions must be considered. The blade boundary layer, influenced byblade’s geometrical parameters and main-flow aerodynamic parameters, usuallydevelops from laminar to turbulence. As a result, to capture the boundary layerdevelopment, and estimating the transition position is significant for theinvestigation of flow performances in the compressor. Besides, variations ofblade profile, blade stacking and end-wall effects often lead to 3Dcharacteristics in the flow fields, where secondary flows, separated flows andcomplicated vortex structures exist. Indeed, it is important to obtain accurateflow information and aerodynamic performances during the design process, whichis crucial for supersonic compressor.

1.     COMPRESSORDESIGN METHODThe stepsinvolved in design of axial flow compressor is as shown in Figure 3 6:By assuming values for the blade tip speed, and theaxial velocity and hub-tip ratio at inlet to the first stage, the requiredannulus area at entry is obtained from the specified mass flow, assumed axialvelocity, and ambient conditions, using the continuity equation.                           Tosatisfy continuity equation:Fromthe above equation,Where is considered as hub to tip ratio and densityis calculated from the relation                                             ……..

(3)                         …ficient in bothmodesgn phase of compressor.ip of rotor and stator blades.peed()be determined.

ThesFrom the suitable design point under sea level conditions(atmospheric pressure and temperature) and from cycle calculations, therequired pressure ration, mass flow rate and compressor inlet temperature canbe determined. These can be useful to calculate the stagnation and staticpressure and temperature at the first stage of compressor.For the assumed axial velocity, tip radius will be the functionof hub to rip ratio.Bladespeed can be represent as:                                      ……. (4)                                       …….

(5)Therefore, we can iteratively determine the tip radius byappropriate hub to tip ratio  and rotational speed (n). It is useful to knowthe mean radius as,Inmean line design methodology, mean radius remain constant for all stages.In case of exit area: Blade height:The radius at exit (final stage) of compressor asdetermined as:                             ……. (9)                                …….. (10)The number of stages is found by dividing total temperaturerise in all stages by temperature rise per stage.Thetemperature rise for a stage is:Thewhirl components of velocity are determined from the velocity triangles asshown in Figure 4.

                The chord length of the blade airfoil is calculatedby appropriate selection of aspect ratio  based on the application.Thepitch and number of blades for the rotor are determined by the given relations,                                             ……(12) 4.1  THE APPLICATION OF SIMILARITYPRINCIPLE:The principle ofsimilarity is a consequence of nature for any physical phenomenon. By makinguse of this principle, it becomes possible to predict the performance of onemachine from the results of tests on a geometrically similar machine, and alsoto predict the performance of the same machine under conditions different fromthe test conditions. If two machines are kinematically similar, the velocityvector diagrams at inlet and outlet of the rotor of one machine must be similarto those of the other. Geometrical similarity of the inlet and outlet velocitydiagrams is, therefore, a necessary condition for dynamic similarity 10.VCE can combine the advantages of turbojet andturbofan by adjusting its bypass ratio in order to change the Thermodynamiccycle parameters.

When the HP compressor was design, two working conditionshave been taken into account, which one is single bypass mode and the other isdouble bypass mode. It means that HP compressor must adapt the bypass ratiovariations which makes the flow rate and compression ratio vary within widerrange than ones of conventional compressor besides the compressor must maintainhigh efficiency. Namely, the flow fields in compressor should be reasonablewhen the cycle parameters are changed greatly 7.

The outer bypass compress air flows into the innerbypass when VCE changes its working conditions. Then the mass flow of HP compressorincreases. The change rate of mass flow depends upon the change of the bypassratio.

There are some different working conditions for VCE matching withdifferent bypass ratios.Assume that theVCE works in n conditions. Then HP compressor must correspond with n seriesof parameters as follows:Condition 1    Condition 2 …………..Condition i ………….

.Condition n Obviously it is different from the conventionalcompressor which is often given only one set of design parameters. It is verydifficult to design the compressor to meet so many series of cycle parameters.Forturbo machinery, the total-total efficiency is commonly depended on tenparameters, namely,    …..

. (14)Thegeometry and gas parameters can be omitted for the same compressor and the sameworking fluid. The following equations can be gotten from the similarityprinciple in turbo machinery.                ……… (15)Where,Compression ratio          Reduced mass flow rate Reynolds number Reduced power Reduced wheel speed Both and  are the qualitativeparameters and can be represented into:            …… (16)               …… (17)Thus,the efficiency is also represented into:                         …… (18)The effect of  can be ignored if Re isbigger than the second critical Re number in turbo machinery. Equation (18) isreduced as:                                                                   …………..

. (19)Therefore, if we make following equations both (20)and (21) to be established for HP compressor of VCE, the multi-conditionscompressor for VCE can be treated as the conventional compressor to bedesigned.As a result, the mass flow rate and inlet parametersof HP compressor are changed into any kind of condition among n seriesof conditions when VCE adjusts its bypass ratio but the variations cannot influencethe efficiency of compressor, which is only influenced by the compressionratio.

2.      ANALYSIS OF DESIGN USING CFD TOOLANSYS Fluent 16.0, a CFD tool has been employedwithin the computations to simulate the 3D steady flows and to validate theaerodynamic performances of designed HP compressor for VCE. Both operatingconditions of single bypass mode and double bypass mode have been taken intoaccount.Analysisof design involves following steps:·        Creating a geometry/mesh·        Defining the physics of the model·        Solving the CFD problem·        Visualizing the results in postprocessor5.1  CREATING A GEOMETRY/MESHThe traditional approach to axial-flow compressoraerodynamic design was to use various families of airfoil as the basis forblade design. American practice was based on various families designed by theNational Advisory Committee for Aeronautics (NACA), the most popular being the65-series family 8.

NACA 65410 9 Airfoil is used here to generate bladecoordinates as in Figure 5.            To create the geometry NX UNIGRAPHICS 9.0 softwareis used. NACA 65410 airfoil coordinates are imported to software and thisairfoil is used from entire hub to tip of rotor and stator blades. Meshing (Figure6) has been carried out by Hyper Mesh. The type of element selected here isTetrahedral.         5.

2  DEFINING THE PHYSICS OF THE MODELIn this step we are defining the physics of Model.This includes specification of type of fluid, defining the domains, Inlet andOutlet Boundary conditions, type of analysis, turbulence model, and heattransfer model etc. The following assumptions are taken for defining physics.·        Steady state condition·        No leakage losses·        Friction between walls and fluid isneglected.5.3  SOLVING THE CFD PROBLEMThesolver parameters are specified as follows:·        Air as an Ideal gas is taken as Workingfluid.·        The one-equation turbulence model ofSpalart-Allmaras has been applied to solve Reynolds’s averaged N-S equations,and this model is applicable to the separation flow simulations of viscousfluid under high pressure gradient.·        For the two operating modes, totaltemperature, total pressure and turbulent viscosity are prescribed at inlet;the gas enters the calculation domain in axial direction.

Meanwhile, staticpressure is given at outlet as shown in Figure 7.·        Two domain interfaces are used. Rotordomain is rotating and stator domain is stationary (Figure 7).Insolver manger the number of iterations and accuracy is specified and results canbe analysed with help of CFD Post.          5.4  RESULTS OF CFD ANALYSIS·        Static pressure distribution on bladesin two modes·        Mach number distribution on the blades·        Vortex movement·        Pressure coefficient distribution onblades in two modes.

6      RESULTANALYSISThe results show that the relative errors of massflow compared to design value. In generally, the results of simulation indicatethat satisfactory aerodynamic performances are available in both of the twooperating conditions, and the designed HP compressor has achieved the anticipatedtargets.                       From the distribution of static pressure on bladesurface (not shown here), the fluid flows around the vane blade characterizedby pressure fluctuating along the whole blade height, and it is the most severenear the hub, which is relevant to wave structures in the channel andcompressive wave boundary layer interactions. These influences are inevitablein the supersonic or transonic compressor.

In the stator, static pressure isfully developed on the pressure surface but in the rotor, static pressure onthe blade presents certain differences in the two modes. It shows the powerrequirement variation in two modes of operation.Isolines distributions of absolute Mach number atstator outlet in each mode represents the characteristics of supersonic wakeflow as depicted in Figure 8. The high-temperature gas flows through thechannel characterized as uneven at different blade heights because of bladestacking and end wall effect. Mixing flow of supersonic fluid leads more energyloss. Moreover, the uneven of flow parameters in the flow fields causes radialmovement from blade bottom to the top and corner vortex (Figure 9) near bothhub and shroud at outlet with secondary flow loss increasing apparently. Thesesecondary flow is mainly responsible for the aerodynamic loss in a supersoniccompressor.         Tip clearance flow is to blame for the aerodynamicloss in both rotor and stator stage, with the proportion to be 45% and 30% insome working conditions.

It is significant to understand the features ofclearance flow for the design of rotor and to minimize the flow loss, sorepresentation formation of tip clearance flow in the supersonic compressor is analysedwith the aid of 3D streamlines as presented in Figures 10a-10c.         There are three forms of tip vortex, (1) Vortexforms behind the guide vane, (2) Vortex forms due to the gas entering the topgap from suction surface relative to high-speed rotating blades, (3) Vortexformed by pressure gradient from blade pressure surface to suction surfacebecause of more tip clearance. Among these three, mass flow rate of the thirdtype of clearance flow is larger than the two previous kinds, and corner vortexoccurs when air current flows out of the suction side.

So it is necessary tocontrol this vortex in design phase of compressor.                         Wave structures of supersonic compressor areanalysed by the variation of absolute Mach number and pressure coefficient atone particular blade height (Figure 11). Detachedshock forms when the gas flow out the trailing edge along pressure surface.Intersection on different side occurs among detached shock, primary expansivewaves and reflective expansive waves. By the variations of Mach number andpressure coefficient, it can be drawn the conclusion that the detached shock isreduced and dissipated by the set of expansive waves.

So, influence of thedetached shock is limited to the flow fields near the trailing edge, while theexpansive wave can occupy most space in the tapered-cutregion.         (b)   Double bypass mode   Figure 11: Mach number and pressure coefficient in both modes               7      CONCLUSIONIn present paper, ahigh loaded high-pressure compressor stage with high compression ratio has beendesigned by NACA profile based on similarity principle. Three-dimensionalcalculations have been carried out under different thermodynamic cycleparameters to simulate flow fields in the two operating conditions (Singlebypass mode and double bypass mode). It reveals the characteristics of supersonicwake flow in a


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