Simulation and experiment of the aerodynamic performance of a vane compressor damaged by a bird strike

Based on the purpose of studying the impact of bird strike on compressor aerodynamic performance, this paper takes a single-stage low-speed axial flow compressor as an object to establish a blade model damaged by bird strike, and on this basis, it carries out a three-dimensional full-loop numerical simulation and experiments on the compressor model with damaged blades. Through comparison with the experimental data of undamaged compressors, the results show that after the rotor blades are damaged, the rotor blades are damaged. The aerodynamic performance and stability of the compressor have significantly decreased, which accords with our understanding. The reliability of the numerical simulation method is verified by analyzing the detailed flow field structure of typical working conditions and comparing the experimental data of the compressor, which lays a foundation for the relevant calculation of damaged blades in the future.


Introduction
As the power source of aircraft, the stability and safety of aero-engine operation are crucial [1] .However, during operation, the engine is likely to inhale various foreign objects.Relevant data show that bird suction accidents are most common, especially after bird impact [2] .After birds strike the blades, the engine blades will be broken and damaged, resulting in large deformation, local depression or leading edge backrolling, resulting in thrust reduction, engine rotor imbalance, deterioration of aerodynamic performance, and even the broken blades will injure other blades, fly out of or penetrate the casing, affecting the normal operation of other aircraft systems and causing relatively serious consequences [3] .As mentioned above, it is of great significance to conduct experimental research on aero-engine bird strike [4] .To avoid the occurrence of bird strike accidents and reduce the economic losses and casualties caused by bird strikes, scholars from various countries have conducted extensive research [5] .In general, the current research on compressor aerodynamic performance of damaged fan blades is very limited [6] .The known foreign research methods are mainly numerical simulation, while the relevant domestic research is in the initial stage [7] .The research on the aerodynamic performance of fan blades after bird strike damage is not perfect or in-depth.Although it is common to land and stop in a short period after a bird strike [8] , the impact of engine blade damage caused by a bird strike on the working characteristics of the compressor is still our concern.Studying the changes in the aerodynamic performance of the compressor caused by bird strikes will help to investigate the aerodynamic stability of the compressor in a more comprehensive way [9] .It provides a reference for us to evaluate the anti-bird strike ability of compressors [10] .This paper takes an axial compressor test bench as the research object to measure and analyze the aerodynamic performance of the compressor.Based on the prototype blade, a model of the blade damaged by a bird strike is established, and a compressor calculation model including the damaged blade is established.Based on this model, the numerical calculation of the aerodynamic performance of the compressor is carried out, and the changes in the aerodynamic performance of the compressor after the damaged blade are compared and analyzed.The reliability of the numerical calculation results was investigated in an attempt to summarize a set of mature experimental and numerical simulation methods for the bird strike phenomenon, which could provide a reference for future research on compressors with damaged blades.

Compressor main design parameters
The schematic diagram of the single-stage axial flow compressor test platform is shown in Figure 1, and the physical diagram is shown in Figure 2. From left to right, there are horns, compressor outer casing, rectifier cone, rotor blade, stator blade, compressor inner casing, support plate, throttle, rotating shaft, speed and torque sensor, and motor.The main pneumatic design parameters of the compressor are shown in Table 1.

Bird strike damage leaf model
In the calculation model used in this paper, a gelatin cylinder with the same compressibility as the mass density of the bird body is used to replicate the load generated by the bird impact.The prototype rotor blade is introduced and divided into hexahedral grids with a mesh size of 3 mm.SPH particles are used in the bird body to impact the upper part of the leading edge of the blade [11] .The finite element model is shown in Figure 3. Bird impact blade finite element model.The blade rotation speed is 6000 rpm and the bird body mass is 100 g.At a speed of 80 m/s, the blade is impacted to the upper part of the leading edge of the blade.Erosion contact is adopted between the blade and the bird body particles.As can be seen from Figure 4, after the impact is completed, the blade does not break, but it is bulging, and the maximum deformation position is about 65% of the height of the leaf.As can be seen from the above figure, although the blade has some deformation and deformation, it remains intact, and the final residual deformation state displacement is mainly the tip displacement and edge bulge.At this point, the bird strike damaged leaf model has been established, which will be used as part of the damaged leaf in the whole loop numerical calculation model in the following paper.

Computational models and meshing
To simulate the bird strike phenomenon, a compressor with three damaged blades uniformly distributed around 120 degrees is used as a model for experimental and numerical calculation.The grid-independent performance of the prototype rotor was verified.When the thickness of the first layer of mesh was 0.007 mm and the total ring mesh size was 1179 w, the grid-independent requirements were met.
The inlet section of the compressor calculation grid is about 2 times the chord edge of the rotor leading edge, and the outlet section is about 5 times the chord edge of the rotor trailing edge.The calculation grid is divided by using the IGG-Autogrid5 module of NUMECA software.The blade area IOP Publishing doi:10.1088/1742-6596/2764/1/0120044 grid is divided by O4H topology, that is, the O-grid is used around the blade, and the H-grid is used in the main flow area and inlet area.Figure 5 is the comparison of the calculated mesh of the rotor blades with 65% blade height B2B surface, and Figure 6 is the calculated mesh of the full-ring blades with damaged blades.

Prototype blade
Damaged blade

Calculation method
The Fine/Turbo module of NUMECA software is used to calculate the steady flow field of the compressor.The governing equations were solved and iterated using the Reynolds mean N-S equation, the fourth-order Runge-Kutta method.The second-order precision central difference scheme was discredited in space.The multi-grid method and the local time step method were used to accelerate convergence, and the turbulent flow model was selected Spalart-Allmaras (S-A) model suitable for turnarachiners.In the calculation, the way of gradually increasing back pressure is used to approach the stall boundary to make the flow transition from the high flow condition to the low flow condition, and the divergence point of the numerical calculation is taken as the stall point of the compressor.

Comparison between prototype compressor experiment and simulation
Figure 7 shows the comparison of experimental and simulated characteristic lines and pressure rise characteristic lines of the prototype compressor at different speeds.The experimental speed in the figure is the reduced speed.As can be seen from the comparison results of experiment and simulation in the figure, the experimental pressure ratio characteristics are consistent with the change law of simulation, but the measured value is generally low, which is because the reduced speed under the physical speed of the experiment is slightly lower than that under the simulation.The difference between the measured value and the calculated value of the pressure ratio begins to increase with the increase of the flow rate, among which the maximum difference between the simulation and experimental results at 20%~100% speed is 0.014%, 0.04%, 0.21%, 0.38%, and 0.59% respectively at the same flow rate.Within the difference range of the reduced speed, the error between the experimentally measured value and the calculated value is acceptable and meets the expected standard.Compressor characteristic line and pressure rise characteristic line.It can be seen from the total static pressure rise characteristic diagram that the variation law of experimental and simulated pressure rise characteristics is consistent at each speed.However, under the same flow coefficient, the pressure rise characteristic line of numerical simulation at different speeds increases with the higher speed.However, for the pressure rise characteristics at different speeds in the experiment, the deviation between each speed is small, and the diagram lines are approximately overlapping.However, it can be seen that the variation rule is the same as the numerical simulation, and the flow coefficient of the stall boundary point and blockage boundary point between each speed are approximately equal.

Anm malysis of aerodynamic characteristics after damage
Figures 8, 9, and 10, respectively show the comparison of experimental and simulation characteristics of the compressor rotor before and after damage at different speeds.It can be seen that the aerodynamic performance of the compressor decreases significantly after the bird strike damage.Taking the design speed as an example, the comparison of numerical calculation results before and after the damage shows that the flow rate at the plug point of the compressor after damage is reduced by about 0.0851 kg/s compared with that without damage, which is a relative decrease of about 0.85%.The flow rate at the unstable boundary point increased by about 0.1624 kg/s compared with that without damage, and the relative increase was about 2.45%.The maximum efficiency is about 0.031 less than that without damage, and the relative reduction is about 3.47%.
At the same time, it can be seen from the comparison of the aerodynamic characteristics of the rotor obtained by numerical simulation before and after damage that although the aerodynamic performance of the rotor after damage is significantly reduced compared with that without damage, the characteristic changes before and after damage are consistent, indicating that the numerical simulation results can reasonably reflect the general law of the influence of the aerodynamic performance of the compressor with damaged blades.Since the back pressure and flow rate of maximum efficiency condition and near-instability condition after blade damage are different, to ensure the comparability of results, the flow field under similar flow rates is compared in this paper.The undamaged compressor is named Model A, and the one with damaged blades is named Model B.

Flow field analysis under maximum efficiency conditions
To analyze the reasons for the deterioration of compressor rotor aerodynamic performance under the condition of maximum efficiency, Figure 11 and Figure 12 respectively give the cloud map distribution of model A and model B relative to Mach number at different blade heights sections.At the 50% blade height section, it is not difficult to see that the flow in the channel changes to some extent due to blade damage and deformation under the maximum efficiency condition, which is significantly different from other channels.When the leading edge of the blade is damaged, the geometric inlet angle of the corresponding blade is increased, and the inflow attack angle is also increased, so there is a low-speed separation area on the suction surface of the damaged blade.However, the flow separation of 90% high section near the blade tip is more serious under the action of tip clearance.The low-speed zone developed from the leading edge of the blade has filled the entire flow channel through the development of the suction surface, which indicates that the low-speed zone formed by the change of the angle of attack of the damaged blade has squeezed the incoming flow of the prototype blade and increased the angle of attack of the incoming flow.However, the low-speed zone does not cross the tip clearance and affects the adjacent undamaged blade passage.Figure 13 shows the relative Mach number cloud image of 50% chord length section of rotor blades in the axial direction of model A and model B. It can be seen that the low-speed zone on the surface of the damaged blade occupies part of the channel area, extending from the damaged location to near the tip of the blade, but the flow of adjacent channels is not affected after being squeezed by the mainstream zone in the channel.At the same time, the overall relative Mach number gradient of the blade surface changes little under the maximum efficiency condition, and no large flow separation is generated, which is also the main reason for low loss and high efficiency under this condition.

Analysis of flow field under near instability condition
Further, the reasons for the deterioration of the aerodynamic performance of the compressor after rotor blade damage under near-instability conditions are analyzed.Figure 14  Figure 16 shows the comparison of the relative Mach number cloud image of Model A and Model B on the section of 50% axial length in the near instability condition.Compared with the maximum efficiency condition, the circumferential range occupied by the low-speed zone in the blade channel composed of damaged blades gradually expands, and the adjacent channels along the rotation direction are affected.It is not difficult to find that if the back pressure continues to rise, the circumferential range occupied by the low-speed zone will continue to expand, making the aerodynamic performance of the compressor continue to decline, and eventually cross the stability boundary and stall.

Conclusion
In this paper, a low-speed axial flow compressor is taken as the object, and a bird strike-damaged blade model is established.The full-ring three-dimensional flow field numerical simulation is carried out for the compressor model with a damaged blade, and the aerodynamic performance and detailed flow field parameters of the compressor blade after damage are obtained.Moreover, the flow field structure is analyzed in detail for two different typical working conditions, maximum efficiency working conditions, and near instability working conditions.By comparing the experimental data of the compressor before and after the damage and the flow field structure under two typical working conditions, it is shown that the full three-dimensional numerical simulation method adopted in this paper for the compressor with damaged blades is feasible, the results are reasonable, and the reliability of the numerical simulation method is verified.It provides technical support for the research of numerical simulation method of compressor aerodynamic performance of bird strike damaged blade.

Figure 1 .
Figure 1.Schematic diagram of single-stage compressor test bench.

Figure 2 .
Figure 2. Physical diagram of single-stage compressor test bench.Table 1. Main design parameters of single-stage compressor.

Figure 3 .
Figure 3. Bird impact blade finite element model.The blade rotation speed is 6000 rpm and the bird body mass is 100 g.At a speed of 80 m/s, the blade is impacted to the upper part of the leading edge of the blade.Erosion contact is adopted between the blade and the bird body particles.As can be seen from Figure4, after the impact is completed, the blade does not break, but it is bulging, and the maximum deformation position is about 65% of the height of the leaf.

Figure 4 .
Figure 4. Blade damage condition.As can be seen from the above figure, although the blade has some deformation and deformation, it remains intact, and the final residual deformation state displacement is mainly the tip displacement and edge bulge.At this point, the bird strike damaged leaf model has been established, which will be used as part of the damaged leaf in the whole loop numerical calculation model in the following paper.

Figure 6 .
Figure 6.Calculation grid of full-ring blade with damaged blade.

Figure 7 .
Figure 7. Compressor characteristic line and pressure rise characteristic line.It can be seen from the total static pressure rise characteristic diagram that the variation law of experimental and simulated pressure rise characteristics is consistent at each speed.However, under the same flow coefficient, the pressure rise characteristic line of numerical simulation at different speeds increases with the higher speed.However, for the pressure rise characteristics at different speeds in the experiment, the deviation between each speed is small, and the diagram lines are approximately overlapping.However, it can be seen that the variation rule is the same as the numerical simulation, and the flow coefficient of the stall boundary point and blockage boundary point between each speed are approximately equal.

6 Figure 8 .
Figure 8.Comparison of compressor characteristic lines after 6000 rpm blade damage.

Figure 9 .
Figure 9.Comparison of compressor characteristic lines after 4800 rpm blade damage.

Figure 10 .
Figure 10.Comparison of compressor characteristic lines after 3600 rpm blade damage.To further analyze the reliability of compressor numerical calculation results after rotor blades are damaged, the flow field characteristics of the compressor are analyzed at different working conditions.

Figure 11 .Figure 12 .
Figure 11.Comparison of relative Mach number cloud image of 50% blade height section (maximum efficiency condition).

Figure 13 .
Figure 13.Relative Mach number comparison of the axial 50% chord length section.

Figure 14 .Figure 15 .
Figure 14.Comparison of relative Mach number cloud image of 50% leaf height cross section (near instability condition).

Figure 16 .
Figure 16.Relative mach number comparison of the axial 50% chord length section.