The results of the first booster stage tone noise numerical investigation for a model of low pressure compressor are presented. The investigation was performed using the frequency domain numerical method of multistage turbomachines tone noise simulation, developed in CIAM (Central Institute of Aviation Motors) and implemented in the 3DAS (3 Dimensional Acoustics Solver) in-house solver. The model under consideration included high bypass ratio fan, stator, booster and bypass duct. Calculation was performed at the approach operating conditions. Far field directivities for two tones in the forward hemisphere were obtained. One tone corresponded to the blade passing frequency of the first stage rotor, the other - to the sum of this frequency with the blade passing frequency of the fan. The results of the computation were compared with the experimental data obtained in the CIAM C-3A acoustic test facility. In general satisfactory correspondence between calculation and experiment was obtained.

The Central Institute of Aviation Motors (CIAM) has been engaged in the development of methods and technologies extending the range of stable operation for GTE axial compressors on the basis of systematic experimental and theoretical investigations of processes before and after flow disturbances for many years. The general sources of experimental data were stage models of various types. They are first supersonic stages with 0.3–0.45 hub ratio and subsonic stages with 0.75 hub ratio, as well as high-loaded stages with low aspect ratio. As a result of these investigations, a structural configuration of the casing treatment (CT) was designed to prevent local flow separation on flow passage surfaces of a compressor stage. The CT structure includes the following components: - Slotted spacer installed above the inlet rotor section; - Attached ring covering the slotted spacer. An approximate procedure for selecting the optimal CT geometric parameters and their interrelations was developed for CT designing. Using this procedure, special investigations were completed and detected the CT effects on operation of the axial compressor. These effects are: - Effect of air back and forward leakage through slots between the blade tips and the inlet rotor section; - Effect of stall deceleration in the stage flow passage; - Pulsation damping at the stage tip when flowing around the CT slotted spacer. Based on this methodology, CT prototypes were developed and tested in various single-stage and multi-stage compressors. As an example of CT advantages, we can show test results for a three-stage low-pressure compressor (LPC) designed by CIAM. The LPC in take-off conditions provides the following design parameters: - Pressure ratio: 3.4; - Corrected tip speed: 418 m/s; - Stall margin: 20% … 21% within 0.5–1.0 corrected RPM. According to experimental investigations, the use of CT results in a considerable increase in LPC stall margin without losses in other design parameters. Additionally, the results of 3D viscous flow calculation are shown for compressor performance analysis.

Results of the high bypass ratio (>8) turbojet fan tone noise simulations in the cutback and sideline operating conditions are presented. Calculations are performed using the method of 3D unsteady numerical calculation of interaction between fan rows implemented in the 3DAS (3 Dimensional Acoustics Solver) CIAM in-house solver. The results of computations are compared with the experimental data for the fan under consideration, obtained in the anechoic chamber of CIAM C-3A acoustic test facility. Comparison was performed for first three harmonics of tone noise for each of conditions. It showed satisfactory qualitative and quantitative agreement between the results of the simulation and the experiment.

Present paper contains a method of solution of inverse problem for Navier-Stokes equations for 3D flows without any simplification of the problem statement and applied to design of turbomachinery bladed rows. In the developed method blade surface is impermeable and no-slip or any other boundary condition compatible with Navier-Stokes equations is applied on the blade surface. Solution of inverse problem is determined using moving grid, which is re-generated at each step of time-marching procedure (variation of flow-rate, impulse and energy fluxes due to movement of grid nodes is taken into account). Normal speed of face of grid cell adjacent to blade surface is determined using given static pressure (inverse mode) with the aid of relationships which are the elements of Godunov scheme applied for integration of flow equations.

When developing counter-rotating fans for advanced new-generation aeroengines with unducted blades it is very important to provide high acoustic and aerodynamic characteristics [1]. This paper presents some results of gasdynamic and aeroacoustic optimization of unducted CRF blade profile by using 3D viscous inverse problem. Flow in unducted CRF on the basis of unsteady 3D Navier-Stokes equations is modeled at the 1st stage of designing in order to find the key tonal noise sources. Based on these results, it is found that one of the key tonal noise sources is Rotor 1 - Rotor 2 tip vortices interaction and potential rotor interaction. Then, using 3D solver of the viscous inverse problem, aerodynamic loads are redistributed along R1 and R2 blade height aiming at a decrease in tip vortex intensity and potential rotor interaction with a probable increase in the CRF thrust. To verify the aerodynamic characteristics of the modified CRF, steady flow calculations are carried out with the help of 3D Navier-Stokes equations and “mixing plane” interfaces. To verify the acoustic characteristics of the modified CRF, tonal noise modeling is carried out for original and modified CRFs using aeroacoustic CIAM’s 3DAS solver for solution of unsteady inviscid equations for disturbances. Ffowcs–Williams, Howkings approach is used for acoustic calculations in the far field. The near acoustic field and directivity diagrams in the far field are found. Using 3D inverse problem, the fan tonal noise is decreased by 4 dB for take-0ff and landing with no thrust and efficiency losses.