Nuclear Engineering and Design

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1 Nuclear Engineering and Design 249 (2012) Contents lists available at SciVerse ScienceDirect Nuclear Engineering and Design jo u r n al hom epage : Highly loaded aerodynamic design and three dimensional performance enhancement of a HTGR helium compressor Tingfeng Ke a,, Qun Zheng b a Shanghai Advanced Research Institute, Chinese Academy of Science, No.99, Haike Road, Zhangjiang Hi-Tech Park, PuDong, Shanghai, PR China b Harbin Engineering University, Harbin , China a r t i c l e i n f o Article history: Received 26 March 2011 Received in revised form 20 March 2012 Accepted 23 March 2012 a b s t r a c t A design study of multistage axial helium compressor of a 300 MWe high temperature gas-cooled reactor is presented in this paper. Helium compressor is characterized by shorter blades, narrow flow channels, numerous stages and longer slim rotor, which result in losses due to blade surface and end wall boundary layers growths, secondary flows and clearance leakage flows, any occurrence of flow separation and stage mismatch. Therefore, the purpose of this paper is to improve and optimize the aerodynamic design of helium compressor. The property of helium is different from that of air, so how to choose the design parameters of a helium compressor is discussed first. And then how to shorten the axial length of the helium compressor or how to decrease the number of stages for a certain pressure ratio by increasing the stage loading are investigated. The new highly loaded helium compressor of larger flow coefficient and high reaction is designed and optimized. Three-dimensional flow patterns in a helium stage are simulated with CFD software (NUMECA). Adjusting the position of blade maximum camber deflection position; redistributing radial compression work; 3D blading techniques, such as distribution spanwise relative airfoil thickness, custom tailoring airfoils and bowed stator vane to mitigate end wall boundary layers and corner separation have improved the aerodynamic performance of the first stage of helium compressor Elsevier B.V. All rights reserved. 1. Introduction The high temperature gas-cooled reactor (HTGR) could be next generation of nuclear power system. For HTGR technology, the power conversion unit currently trend towards utilizing gas turbine system because of its higher efficiency and simple structure. For example, the projects of PBMR, GT-MHR, GTHTR300, etc. are designed to utilize gas turbine generator coupled with HTGR (Kostin et al., 2004; Kunitomi et al., 2004; Matzner, 2004). Now there are two small experimental HTRs are presently under development: in Japan, the high temperature test reactor went critical and is going up in power while being tested. Its normal power of 30 MW should be reached now. It used bloc type fuel of Japanese design somewhat different from the American one. In China, HTR 10 is a 10 MW reactor which had gone critical in It was designed with the collaboration of German companies and uses pebble type fuel. The helium compressor plays an important role in the HTGR system as helium at the desired operating conditions as the coolant for Corresponding author. address: ketingfeng313@126.com (T. Ke). the HTGR reactor and acts as the working fluid of the power conversion unit. Design and optimization of a multistage turbine for helium cooled rector has been researched (Van den Braembussche and Brouckaert, 2008). A model, which is able to describe compressor behavior in off-design conditions and especially full stall condition, has been provided for a nuclear gas turbine plant (Tauveron, 2008). Using helium as working fluid, however, several special conditions of both pro and con to the unit aerodynamic design have to be considered. The sonic speed of the light molecule weight helium gas is sufficiently high, about three times of air s, making the helium gas flow at low subsonic and generally free of shocks in the turbine and compressor. The other working conditions are less than favorable either, especially in the compressor whose aerodynamic design targets contemporary performance goals of 90% polytropic efficiency and 20% surge margin in quest of base load efficiency and reliable nuclear operation. The specific heat of helium is five times than that of air, making helium about as many times less compressible than air such that it needs many more stages to compress helium for a given pressure ratio. Furthermore, the volume flow remains small despite the unit s high power rating because the compressor is designed to work at an inlet pressure of 24 times the atmospheric pressure as compared to the usual 1 atm at air gas turbine /$ see front matter 2012 Elsevier B.V. All rights reserved.

2 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Nomenclature CFD computational fluid dynamics CDA controlled diffusion airfoil HTGR high temperature gas cooled reactor PBMR pebble bed modular reactor GT-MHR gas turbine-modular high temperature reactor GTHTR gas turbine high temperature reactor s isentropic efficiency ϕ flow coefficient C u absolute velocity on peripheral velocity r radial axis W u change of tangential gas velocity compressor inlet. These unfavorable conditions lead to what appears to be a large number of blade rows, though pressure ratio is rated low, populating a narrow and essentially parallel compressor flow path. The multistage slender flow path gives rise to the relative importance of aerodynamic losses associated with blade surface and end wall boundary layers growth, secondary and clearance flows, and any occurrence of flow separation and stage mismatch, making overall efficient compressor performance difficult to achieve. The slender compressor rotor will result in dynamic problems (Kim et al., 2008; Muto et al., 1999). The investigation of the highly loaded blade airfoil is a key design issue in the development and deployment of helium compressor for high temperature gas cooled reactor. The analysis and conclusion presented in this paper can be used as a reference to helium compressor design and the experimental research works. 2. Highly loaded design of helium compressor The choice of working fluid affects the turbocompressor primarily in the following two ways: (1) the number of stages for the attainment of the required pressure ratio and efficiency; and (2) the machine size for a high pressure closed loop system. The specific heat of helium is five times that of air, and since the stages temperature rise varies inversely to the specific heat (for a given limiting blade speed), it follows that the temperature rise available per stage when running with helium will be only one-fifth that of air, and this of course, results in more stages being required for a helium compressor. The physical properties of helium and air are given in Table 1. As mentioned previously the required pressure ratio in a highly recuperated closed Brayton cycle results in a number of compressor stages that is comparable with existing air breathing gas turbines (Mcdonald et al., 1994). Substitution of helium for air greatly modifies aerodynamic requirements by removing Mach number limitations; the problem then becomes that of trying to increase the highest possible gas velocities that stress-limited blades will allow. For the selected machine configuration the compressor rotational speed is fixed at 3600 rpm. The size of the machine is thus dictated by the choice of blade speed, there being an incentive to use the highest values commensurate with stress limits to reduce the number of stages, since the stage loading factor is inversely proportional to the square of the blade speed. Helium compressor for a closed cycle gas turbine is characterized by small blade height, high hub-to-tip ratio and low aspect ratio. Thus, careful mechanical design is necessary to minimize tip clearance effects and the end-wall loss. While the end-wall boundary layers have an adverse effect on efficiency, three other factors of helium compressor that will partially offset this factor are: (1) high Reynolds numbers ( ); (2) very low Mach number (<0.40); and (3) retention of clean oxide-free blade surfaces in the inert helium environment. The present base line design approach to high performance helium aerodynamics rules out compressor intercooling because intercooling optimizes cycle performance only at demanding high compression ratio that results in significant machinery complexity and inefficiency in delivering helium. The optimum pressure ratio is greatly reduced in the present non-intercooled compressor that employs 20 stages to deliver a design pressure ratio of 2.0. In addition to shorter bladed flow path, reduced pressure ratio in a given cycle, heat rate boosts volume flow, which can effectively improve the flow area of a generally narrow helium flow path. The blades are allowed to be of sufficient span to keep blade essential core passage from end wall and tip clearance effects. This is done in the present design without resorting to high rotor speed. Instead, the speed is set to permit efficient gas turbine drive of both the compressor and the grid-synchronous generator on single shaft. A second base line design choice made is the use of high reaction blading to balance the dual design goals of efficiency and surge margin achievable in a reasonable number of compressor stages. The classic air compressor practice has preferred use of lower reaction blading. On the other hand, the EVO operating helium unit used a much higher reaction of 100% (Yan et al., 2008). In our following studies, some advanced blading techniques were adopted in helium compressor for the first time. Hubbs and Weingold described the controlled diffusion airfoil CDA approach to air transonic blades (Tauveron, 2008). MHI had used double circular arc rotor blades and NACA65 stator vanes until CDA was developed, which was used in MF221 and M501G combustion gas turbine compressors in the 1990s, and gained efficiency and operating range over the conventional airfoils when given the same vector diagrams. The present helium design uses a type of CDA for subsonic flow with specific design goals, which include adjusting incidence to achieve single stagnation point near the leading edge, limiting peak Mach number on the suction surface, and achieving suction surface diffusion from the peak Mach number until the trailing edge without separation and with the well-known Stratford velocity profile of minimum skin friction. Wisler showed through low speed model testing how custom-tailored airfoils were used to improve a range of performance goals such as reducing separation and end wall losses, which are particularly relevant to helium application (Wisler, 1985). In a similar approach, the present design uses a proper degree of stator bow to take out suction surface corner separation and the 3D stacking of camber and stagger to restore flow incidence near end walls. This paper will try some design approaches to get a highly loaded helium compressor by increasing the inlet axial velocity value greatly. Then improved design has been done for the threedimensional flows in the helium compressor. Table 1 Helium and air physical properties. a Gas C p (kj/(kg K)) K (kg/m 3 ) ( Pa s) R (J/(kg K)) Helium Air a NTP (normal temperature and pressure) is defined at K and 1 atm ( kpa) (Weisbrodt, 1994). 3. The design and internal flow analysis of highly loaded helium compressor 3.1. Highly loaded design Based on the project of Japanese 300 MW low-pressure helium as a baseline design (Muto and Ishiyama, 2000), the present highly

3 258 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Table 2 The comparison of main parameters. Parameters Baseline design Highly loaded design Inlet pressure [Mpa] Inlet temperature [ C] Single stage pressure ratio(1th) Compressor pressure ratio Mass flow [kg/s] Axial velocity[m/s] Number of stages Aerodynamic design pitch line values Tip diameter (first rotor) [m] Hub diameter (first stator) [m] Tip speed (first stage) [m/s] Rotor/stator vane/blade count (first stage) 142/ /199 Rotor/stator chord (first stage) [mm] / / Rotor/stator solidity (first stage) 1.4/ /1.4 Rotor/stator aspect ratio (first stage) 1.7/ / 1.6 Rotor tip/stator hub clearance 1% blade span 1% blade span Flow coefficient Load coefficient Reaction loaded helium compressor is designed and compared with the baseline design as shown in Table 2. As shown in the table, the stage number is reduced from 16 to 10 to shorten the rotor, the flow coefficient and load coefficients are increased, but the aspect ratio is decreased in the highly loaded design. The baseline design of the helium compressor has been considered and designed very well; the flow coefficient of the helium cascade is 0.6, which is under the maximum efficiency of airfoil. The stage loading and reaction must be increased together to maintain optimum efficiency for a nominal stage in a low speed compressor. The choice of stage loading and stage reaction are inextricably linked. They must be considered together to allow stage loading to be successfully advanced beyond currently comfortable levels whatever the design flow coefficient (Dickens and Day, 2009). Otherwise the baseline design also has a high degree of reaction, so the value in highly loaded design was selected as 0.7 maybe some higher value. In order to reduce the helium compressor stages, giving fully using to the characteristics of low Mach number, the Inlet axial velocity can be increased properly, so as to increase the flow coefficient, and also the single stage pressure ratio. This relation of stage efficiency and flow coefficient is different between air and helium. We choose the flow coefficient as 0.96 in our highly loaded design, which efficiency is just a little lower than the maximum value as shown in Fig. 1. This is also the reason why the helium compressor has the high Reynolds numbers. Then three dimensional aerodynamic optimization of the highly loaded designed helium compressor cascade is carried out to improve the highly loaded compressor efficiency. The main difference between helium and air turbomachine is the much larger change in blade height of the later one. The meridional contour of highly loaded air compressor is strongly diverging whereas helium compressors have an almost constant blade height, allowing the use of nearly identical stages from inlet to outlet. This means that the design of a helium compressor can be concentrated on one stage. Main design parameter of this stage (first stage of the helium compressor) was listed in Table 2. The numerical simulation was performed using the commercial CFD software package NUMECA. The Navier-Stokes equations are discretized with a cell-centered explicit finite volume scheme according to Jameson in a relative coordinate system rotating together with the reference frame. Time integration is a four-step Runge-Kutta algorithm. In order to speed up the convergence, local time steps, residual smoothing and multigrid techniques were applied. The Spalart-Allmaras turbulence model was used for the closure of the equations. The computational mesh was generated with grid generation tool of NUMECA/Autogrid/IGG. For fulfilling the requirement of the turbulence model, y+ was controlled within 10. Total grid number of a single blade passage was about 700,000. Fig. 2 shows the medium grids of the computational mesh. The characteristic values are as followings: characteristic density is 3.7 kg/m 3 ; characteristic velocity is m/s; characteristic length is m. Boundary conditions are defined as: inlet total pressure 2,442,000 Pa, inlet total temperature 308 K, and outlet static pressure 2,451, Pa Analyses of the helium compressor flow field Comparison of the static pressure coefficient distributions In the highly loaded design, the camber line and thickness distribution curves of the baseline are both formed by quadratic polynomials. In the design process, controlling the maximum deflection value and the position of the maximum camber in tenths Fig. 1. Relation of stage efficiency and flow coefficient under different reaction degrees for helium.

4 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) making the minimum pressure point located at the position around 12% of the airfoil in percentage of chord, and further by controlling the inverse pressure gradient, which also benefit to suppress the development of the boundary layer. The distribution of static pressure of stator blade surface is more favorable after highly loaded design, as shown in right one. But the steep pressure gradient on the suction surface is still too high, especially near the leading edge. The match of rotor and stator is also important when optimization is applied to improve the profile efficiency. Fig. 2. The computational mesh. of chord, and the maximum thickness of the airfoil in percentage of chord, ensure the turning angle requirements of the design. The static pressure rise coefficient is defined as: p p C p = inlet 0.5 inlet W 2 inlet where p inlet is inlet mass-averaged static pressure, inlet is inlet mass-averaged helium density, W inlet inlet mass-averaged relative velocity. The pressure coefficient distributions on the blade surfaces of the rotor and stator are shown in Fig. 3. The area surrounded by the pressure lines represents the load of the blade. It can be observed that both rotor and stator blade loads have greatly increased as seen in Fig. 3. In left one, the minimum pressure points on pressure side of the highly loaded designed blade surfaces are located very near the leading edge, and the inverse pressure gradient is relatively long, which is no good for the cascades flows and may result in serious profile losses. The airfoil will then be optimized by controlling the lowest pressure point position, moving it backward, Secondary flow characteristics The essential secondary flow phenomenon in the cascade flow is the cross flow on the end-wall caused by the pressure gradient between the pressure side of a blade and the suction side of an adjacent blade. The low momentum fluid is forced towards the suction side corner, where this flow separates due to the high pressure gradient in flow direction. The typical flow pattern for these separations is visible in Fig. 4. The separation influences the deflection at the end-wall region and is responsible for a major part of the losses. Also the cascade performance in the form of flow turning and pressure rise is influenced by the scale and intensity of the separation. On rotor suction surface, as shown in Fig. 4, for the baseline design, the separation occupied about 25% of the blade passage at the near hub region, and it extended about 40% of the blade height in spanwise direction. While in highly loaded design, the separation area became larger, which occupied about 40% of the blade passage at the near hub region, and it extended about 90% of the blade height in spanwise direction. On stator suction surface, the flow separations in the corner region also become more obvious after highly loaded design. This happened due to the increase of stage load Profile loss Distributions of relative Mach number. Fig. 5 plots the distribution of absolute Mach number of the blades. In spite of the increasing of flow coefficient from 0.6 to 0.96 and the increasing of axial velocity simultaneously, the Mach number is relatively low in the whole flow region because of the higher sound speed of the helium. As shown in Fig. 5, for the highly loaded design, the Mach number in the whole flow region is in the range of , little larger than the range of of the prototype. Comparing the distributions of Mach number, the difference of the development C p 1 Baseline design Highly loaded design C p 0.8 Baseline design Highly loaded design X/C Rotor X/C Stator Fig. 3. Distribution of static pressure coefficient at 50% blade height.

5 260 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Three-dimensional improved design The highly loaded design of the helium compressor cascade has been implemented as shown above, the stage load increased, and the stage number is decreased from 16 to 10 stages, but the flow fields are not so good due to the increased stage load. Some optimizations are necessary to improve the performance of the helium compressor. The compressor blade stacking and blade profile is optimized by three dimensional CFD simulations Improved design of the chordwise camber Fig. 4. Distribution of limiting streamlines on the blade surfaces. of boundary layers also can be observed which indicates the profile loss to a certain extent Total pressure loss coefficient. The profile losses can be observed clearly from the distribution of total pressure loss coefficient at blade exit, which includes the cascade total pressure loss and additional loss of total pressure due to the wake fluid turbulent mixing after the trailing edge of the blade. Fig. 6 describes the azimuthally averaged total pressure loss coefficient at rotor and stator exit. The total pressure loss coefficient is defined as: ω = (P 0e P 0i )/P vi, where P 0e is the total pressure at outlet, P 0i is the total pressure at inlet, P vi is the dynamic pressure at cascade inlet. As shown in Fig. 6 the high loss regions both in baseline design and highly loaded design exit at two ends extending about 10% height of blade. At rotor exit of highly loaded design, the total pressure coefficient increase from 15% to 45% height of blade while loss decreased compared with baseline design, and above 55% the total pressure coefficient in highly loaded design greater than in baseline design in contrary, which indicates that the flow field at the upper half of cascade has deteriorated due to highly loaded. To stator cascade, the total pressure coefficient has increased in the main flow field about 20 90% height of vane. Other flow field basically keeps the same with the baseline design. The maximum camber of the general airfoil is adjusted, and the position of this maximum camber is moved and optimized to form a helium-foil. The CFD simulation results are as shown in Table 3. With the position of the maximum camber moving forward, the efficiency and pressure ratio are both increased. As the maximum camber position moves forward, flow rate and pressure ratio increase. There is an optimum value for the efficiency. In Fig. 7 the load on the helium-foil increased with the maximum camber deflection position moving forward and inverse pressure gradient on the pressure surface is not so steep. At the suction surface the pressure gradient increase abruptly from the leading edge to the 30% axial chord length of the blade, and then increase gradually similar to that with the maximum camber position at 0.5 chord length. If the maximum camber position is at 0.5 chord length, the load distribution shown with the solid line is uniform on the helium-foil, but after changing the maximum camber position to 0.3, the highest load moves forward, as shown in the dashed line. For the load on the stator vane, as shown in right one, no notable change was observed when the maximum camber position moving from 0.3 to 0.5, the load at leading edge area increases a little in the case of maximum camber position is 0.3, and the whole static pressure coefficient increase. It is concluded that the change of the maximum camber deflection position can affect the cascade diffusion ability, and the blade surface velocity distribution, accordingly affect the boundary layer thickness, as well as three dimensional flows. For helium compressor, the flow is also in subsonic condition though the load increasing obviously. Moreover the boundary layers on helium blade surface develop rapidly for high kinematics viscosity coefficient of helium. When moving forward the maximum camber deflection position, the flow turning angle will allot at longer diffusion regions, consequently controlling the development of the boundary layers. From the limiting streamlines distributions on the suction surfaces of the rotor, the movements of the maximum camber positions have an effect on the separations on the suction surface. In case 1, the suction surface separation line originates at 55% chord from the leading edge and terminates at the 90% height of blade, while in case 4 the suction surface separation line originates at 40% chord from the leading edge, earlier than case 1, and extend only 15% height of blade. It is indicated the negative pressure gradient slow up in case 4, then thickening the surface boundary layers close to the end-wall, resulting in the diminish of corner separation (Fig. 8) Improved design of spanwise load distribution Fig. 5. Relative Mach number at 50% blade height. The spanwise swirl distribution (sometimes referred to as the vortex design is a major contributor to the overall performance of a multistage machine. Changing the swirl distribution can be considered a 2-D design change and overall effects can be predicted using three dimensional CFD simulation. The designs discussed above, the baseline design and highly loaded design, which were both based on C u r = constant, result in the low energy fluid accumulate on the end-wall corners. In order to decrease the loss and increase the

6 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Fig. 6. Mass-averaged total pressure loss coefficient at rotor and stator exit. Table 3 Computational results with different maximum camber position. Position Efficiency Static pressure ratio Static pressure ratio Isentropic Polytropic 0.5 (case 1) (case 2) (case 3) (case 4) stage efficiency, the following improved design will redistribute radial compression work, which is reducing the ends values of C u r, increasing the middle load, so that the turning angles at the ends, particularly at the root of the blade, were decreased. The blade form and turning angle are adjusted, and the lowest pressure point is controlled at near 20% of the chord. In the following discuss, the highly loaded design with the position of maximum camber 0.3 is defined as Rotor A, and the modern controlled vortex design is defined as Rotor B. Because the swirl distribution changed, the exit swirl angle has changed, so the stacking lines at leading edge and trialing edge become irregular, shown as Fig. 9. Based on the pressure ratio stand, the isentropic efficiency has increased 1 point (Table 4). The separation area at the trialing near hub has controlled in Rotor B. And the low pressure region has covered more regions at the leading of suction surface. We can also find the flow structures in the separated regions are a little different. Refer to the flowing detailed discussions, as shown in Fig. 10. Comparing the detailed flow structures at separated regions of Rotor A and Rotor B, the most different part is that there is a saddlepocus pair on the suction surface, which lies at a distance of about 95% chord from the leading edge and 18% span. That means it is thicker of the separated layer close to the end-wall in Rotor A, 0.2 Case 1 Case 4 C p 0.2 Case 1 Case 4 C p X/C Rotor Stator X/C Fig. 7. Distribution of static pressure coefficient at 50% blade height.

7 262 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Table 4 Computational results with different spanwise load distribution. Static pressure ratio () Pressure ratio stagnation (*) Isentropic efficiency ( s) Rotor A Rotor B Fig. 8. Distribution of limiting streamlines on the rotor blade surfaces. Fig. 11. The total pressure loss coefficient at rotor exit. controlled, which are under greater influence of the adverse pressure gradient. As depicted in Fig. 11, the total pressure loss coefficient distributions have a little discrepancy. At hub, tip and middle sections it is almost the same, while at 20 30% span and 60 80% span, the total pressure loss coefficient has decreased in Rotor B. Fig D rotors. and after improved design, in Rotor B, the flow lifting off at the trialing is diminished a lot. On the end-wall, the trailing saddle point S has emerged a little earlier in Rotor A than in Rotor B predicting that the limiting streamlines undergoing flow reversal has 4.3. Improved design of spanwise distribution of relative airfoil thickness Helium compressors for closed cycle gas turbines are characterized by small blade height, while highly loaded design, the blade height will become shorten, which will decrease the demand for intensity of vibrations. In the flowing study, we try to compare the Fig. 10. The detail flow structures at separated regions.

8 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Fig. 12. Relative thickness changed from hub to shroud. 1 1 h h Rotor B Rotor C Rotor B Rotor C C a /(m/s) C a /(m/s) Fig. 13. Traversed measurements at rotor exit. relative thickness changed from hub to tip to same thickness distribution from hub to tip. It can be directly observed the change of relative thickness in spanwise in Fig. 12; in view of the short blade, the variable amplitude is not very clearly. At the hub section the relative thickness is 10% chord, and then smooth diminish to 8% chord at tip section, comparing the Rotor B relative thickness keeping the same from hub to tip. Based on the pressure ratio stand, the isentropic efficiency has increased 0.4 point, indicating the relative airfoil thickness changing has positive affection to flow field (Table 5). Fig. 13 shows the spanwise traverse measurements of axial and tangential flow velocities in the wake of rotor exits. As shown in Fig. 13, the left one axial velocity distribution in spanwise, from 15% to 35% and 55% to 95% height of blade, the axial velocity at Rotor B exit is little greater than Rotor C; from 35% to 55% height of blade two cases has the same tendency; below 15% height, near hub endwall, Rotor B exhibits a little bigger axial velocity. It can be explained that the relative airfoil thickness decrease from hub to tip section, meaning the cascade diffusion area increase, so the axial velocity will increase with the same flow rate. The simulation value of Rotor C is more close to the average axial velocity 240 m/s in 2-D preliminary design. A pike in the tangential velocity profile appears near the casing wall. It comes from interaction of tip clearance leakage and vortex with the casing boundary layer. The Table 5 Computational results with different airfoil thickness. Pressure ratio stagnation Isentropic efficiency Rotor B Rotor C spike is less visible in Rotor B maybe because the wall boundary layer there remains undeveloped Custom tailoring airfoils As Mach numbers and aerodynamic loading on airfoils increased and as improved cascade analysis techniques became available, designers recognized that the standard airfoil types conventionally used could often be improved. Consequently, in an effort to reduce loss, reduce flow separation, reduce the effects of tip leakage and secondary flow, control boundary layer growth, etc. in the following discussion, the camber and thickness distribution will be tailored to improve the stage efficiency. The meanline angle distribution and the airfoil section shape for the hub of the Rotor C are shown in Fig. 14. A subsequent custom tailoring of the airfoil section near the rotor hub section by uncambering (unloading or straightening) the trailing edge region, overcambering (loading) the leading edge region to control the separated flow on rotor hub and boundary layers on endwall. Consequently, the tip section meanline shape was modified to loaded trailing edge and unloaded leading edge. As shown in Fig. 14 dashed lines, comparison Rotor C hub meanline with custom-tailored Rotor D, the meanline shape by adding curve slop gradually near leading edge region is able to realize the flow rapidly diffuse at leading edge region and then bring down the flow velocity. The flow turning angles are great with steady changing rate at most middle regions of blade surface. And for trailing edge region, the turning slop of mealine curve decrease gradually by flattening trailing edge region, which is benefit to reduce the boundary layer on suction surface derived from adverse pressure gradient and reduce corner separated region. Fig. 15 describes the static pressure

9 264 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Fig. 14. Comparison of the Rotor C hub section with custom-tailored Rotor D to reduce separation. Fig. 16. Comparison of the Rotor C tip section with custom-tailored Rotor D to reduce tip leakage. coefficient on hub section blade surface, as shown in figure, the static pressure rise curve character by more smooth and pressure rise being controlled at the rest half of suction surface, which indices the leading region has achieved most of the pressure diffusion, thereby eliminate the surface boundary layers. Since the flow velocity keeps high on forward portion of airfoil and the boundary layers are stable and capable of large amounts of diffusion without separation. As the turbulent boundary layer developed, the diffusion rate is decreased toward the trailing edge. The comparison of the tip section mealine and airfoil type is given in Fig. 16. It was thought that the tip clearance flow might be entraining suction surface boundary layer fluid, explaining the unusually thin wakes near the tip. The rotor trailing edge should, therefore, be able to take higher diffusion rates without separation. Consequently, the tip meanline shape was modified to unload the leading edge by removing 4 5 deg of camber and loading the trailing edge by adding 6.0 deg of camber relative to Rotor D. the maximum thickness was moved from 45 to 70 percent chord, which gave increased trailing edge loading. This could reduce the maximum pressure difference across the rotor without undue risk of f low separation. Also measurement blade surface static pressure coefficient distribution showed that the increased trailing edge loading and reduced peak-suction-surface velocity were achieved (Fig. 17). When given the cascade inlet flow angle and cascade metal outlet angle, the spanwise distribution of outlet flow angle indicate the blade surface boundary layers development, and distribution of load and diffusion ability in spanwise. The absolute flow angle at rotor exit was shown in Fig. 18. The most obvious trend is the reduction in deviation angle, which results in the substantial increase flow and pressure rise and removal of flow separation. From 70 to 90 percent of blade height the turning angle has increased Improved design of bowed stator vane Fig. 15. Elimination of flow separation by measurements of blade surface static pressure coefficient. Numerous investigation showed that the bowed blade with suitable lean length and lean angle would reduce the endwall losses effectively, Using bowed blade could avoid the flow separation in the comers of the suction surface and endwalls, while increasing

10 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Fig. 20. Stator vane structure. Fig. 17. Blade surface static pressure coefficient. Fig. 21. Streamline distribution on suction surfaces. Fig. 18. Spanwise absolute flow angle at rotor exit. the compressor efficiency. So the following improvement will try to improve end-wall flow at stator vane with bowed stator. All above improved designs are concluded in Table 6. The characteristic features of a bowed vane are a positive lean angle 5 at the hub with 10% height vane lean length and a negative lean angle 1 at the shroud with 90% height vane lean length in comparison with a normal, radially stacked vane. The detail bowed blade define method refers to Fig. 19. To study the effects of bowed Fig. 19. Bowed blade design method. stator vanes on compressor performance and on the radial distribution of aerodynamic loading, there is no change of the rotor blading (Rotor D) and the vane count. The comparison of stator vans are given at Fig. 20. The limiting streamlines on the suction surface are presented in Fig. 21. The flow separation near casing was eliminated obviously. The features present in the exit flowfield of a blade row, such as wakes and streamwise vortices, will eventually mix out and so generate losses (Denton, 1993). The amount of entropy produced during this process is dependent upon the environment in which the mixing takes place. For compressors, it has been found that the stretching of the wake within the next blade row reduces its velocity deficit, so that the loss generated by mixing the wake within the row is less than if it had mixed out upstream of the row (Smith, 1966; Valkov and Tan, 1999). The comparison of secondary flow velocity vectors at stator exits is shown in Fig. 22, from which we can see clearly the strength and location of the passage vortices. A fully developed passage vortex was shown at both upper end-wall sides, which is close to the suction surface at this location and the core is removed from the pressure surface to suction side. From the velocity vectors it is evident that the passage vortex occupies a much larger region of the endwall in Stator I for the thicker boundary layer. Bowed stator has reduced the centrifugal effects resulting in the reduction of an accumulation of low energy boundary layers. It can be seen that in Stator I, the passage vortex core is an unstable spiral node, while in Stator II, it changes to a stable spiral node, that means the streamline near the vortex is pointed inward towards the core. At lower end-wall, the pressure gradient has also reduced. So by using bowed blades it is possible to strengthen the stability of the passage vortex.

11 266 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) Table 6 Listing of blading designs presented in text. Airfoil Type Rotor Stator Case 1 Stator I Maximum camber position: 0.5 chord, NACA thickness distribution Case 2 Maximum camber position: 0.45 chord, NACA thickness distribution Case 3 Maximum camber position: 0.4 chord, NACA thickness distribution Case 4 = Rotor A Maximum camber position: 0.3 chord, NACA thickness distribution Rotor B Redistribute spanwise compression work Rotor C Redistribution spanwise relative foil thickness Rotor D Custom tailor hub and shroud sections by using camber and thickness distribution Rotor D + Stator II Bowed stator Fig. 23 summarizes the total pressure loss coefficient of all the blading improved design. The loss has decreased from Rotor A to Rotor D + Stator II , nearly 2 point. 5. Overall performance The single stage off-design performance at design rotational speed is obtained by analyzing the compressor stage at different pressure ratios. This induces a change in mass flow and hence of the compressor exit flow conditions that are the next stage inlet conditions. As shown in Fig. 24, the peak efficiency flow rate and the shape of the characteristic near stall are well predicted. The pressure ratio and isentropic efficiency are both increased, comparing Rotor D + Stator I and Rotor D + Stator II, indicating bowed stator has increase the cascade diffusion capacity and also the stable operating range. Fig. 22. Distribution of secondary flow velocity vectors. Fig. 23. Total pressure loss coefficient summarizing with different blading designs (rotor exit). Fig. 24. Performance curves at design speed (3600 rpm).

12 T. Ke, Q. Zheng / Nuclear Engineering and Design 249 (2012) techniques have adopted to mitigate end wall boundary layers and corner separation. Change the position of maximum camber deflection position. As the maximum camber position moves forward, flow rate and pressure ratio have increased and three dimensional separations diminished a lot due to the optimization of blade surface pressure gradient. Redistribute the load in spanwise (such as reducing the flow angle in hub region, and increasing the work input in rotor middle region). The separation area at the trialing near hub has been controlled. Improve design of spanwise distribution of relative airfoil thickness which optimizes the cascade diffusion flow field resulted in stage efficiency increased 0.4 point. Custom tailoring hub and tip airfoils. It has a positive effect to reduce loss, reduce flow separation, and reduce the effects of tip leakage and secondary flow, control boundary layer growth by adjusting camber and thickness distribution. It is found that bowed stator vane in helium compressor can also strengthen the stability of the passage vortex, and then reduce the passage vortex losses. And even increase the performance and also the stable operating range at design speed. Comparing the performance curves at design speed, the final highly loaded design Rotor D + Stator II has increased the pressure ratio obviously, and keep the stage efficient high, even with large stable operating range. Acknowledgement This work is supported by the National Natural Science Foundation of China under Grant No References Fig. 25. Performance curves of final design at different speeds. Comparing the baseline and final highly loaded design Rotor D + Stator II, the stage pressure ratio has increased obviously, and keep the stage efficient high, even 2 points larger than baseline at design point. With the flow rate reduced, the efficiency increased in baseline more rapidly than highly loaded design until near the stall angle of attack is reached. It is shown that the stable operating range has not diminished though the stage loading enhanced greatly. And the performance curves of final design at different speeds are also reasonable (Fig. 25). 6. Conclusions A highly loaded helium compressor is designed and analyzed in this paper. The basic and advanced approaches are taken in the present aerodynamic design for the multistage axial helium compressor of a 300 MW class nuclear gas turbine. The large flow coefficient and load coefficient are selected to provide the flow path working conditions such as pressure ratio and volume flow that result in the flow path configuration of a single bladed rotor with reasonable stage count and sufficient flow area. The selection of the high reaction blading is intended to obtain dual design goals of efficiency and operating stability. The advanced 3D blading Denton, J., Loss mechanisms in turbomachines. ASME J. Turbomach. 115, Dickens, T., Day, I., The design of highly loaded axial compressors. ASME. Kim, J.H., No, H.C., Kim, H.M., Direct implementation of an axial-flow helium gas turbine tool in a system analysis tool for HTGRs. Nucl. Eng. Des. Kostin, V.I., Kodochigov, N.G., et al., Power conversion unit with direct gasturbine cycle for electric power generation as a part of GT-MHR Reactor Plant. In: Proceedings of the 2nd International Topical Meeting on High Temperature Reactor Technology, Beijing, China. Kunitomi, K., Yan, X., et al., GTHTR300C for hydrogen cogeneration. In: Proceedings of the 2nd International Topical Meeting on High Temperature Reactor Technology, Beijing, China. Matzner, D., PBMR project status and the way ahead. In: Proceedings of the 2nd International Topical Meeting on High Temperature Reactor Technology, Beijing, China. Mcdonald, C.F., Orlando, R.J., Cotzas, G.M., Helium turbomachine design for GT-MHR power plant. ASME. Muto, Y., Ishiyama, S., Design study of helium turbine for the 300MW HTGR-GT power plant. ASME, Muto, Y., et al., Design study of helium turbine for the 600MWt HTGR-GT power plant. In: International Gas Turbine Congress, Kobe. Smith, L.H., Wake dispersion in turbomachines. ASME J. Basic Eng. 88, Tauveron, N., Analysis of stalled compressor and application to a nuclear gas turbine plant. ASME, June Valkov, T.V., Tan, C.S., Effect of upstream rotor vortical distribances on the time-averaged performance of axial compressor stators. Part I. ASME J. Turbomach. 121, Van den Braembussche, R.A., Brouckaert, J.F., Design and optimization of a multistage turbine for helium cooled reactor. Nucl. Eng. Des. Weisbrodt, I.A., Summary Report on Technical Experiences from High Temperature Helium Turbomachinery Testing in Germany. IAEA, Vienna, Austria. Wisler, D.C., Loss reduction in axial-flow compressors through low-speed model testing. ASME J. Eng. Gas Turbines Power 107, Yan, X., Takizuka, T., Kunitomi, K., Aerodynamic design, model test, and CFD analysis for multistage axial helium compressor. J. Turbomach.

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