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.

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.

This work presents the latest results of aeromechanical design of two large-scale fan model stages (Dr = 700 mm) for low-noise high-performance single-stage fan prototypes designed for advanced civil aircraft geared and direct-driven turbofans with reduced and ultra-low rotor tip speeds, high specific capacity, and high bypass ratios. They are designed with account of all features of blades made of polymer composite materials (PCM) or titanium alloy. Metal and composite blades have a similar shape in hot state at the design point. The stages are intended for tests in the anechoic chamber of the CIAM’s C-3A special acoustic test facility with the aim of verification new optimal design methods for similar fans to achieve maximum performance. Performances of the fans and parameters of viscous steady flows are calculated. The calculations show that both fan models can provide a high specific capacity along with a high efficiency and sufficient stall margins. For example, calculated max. efficiency level of the bypass duct in the geared model fan with ultra-low tip speed (Ucor. = 313.4 m/s) is equal to 94%. Data measured by tests of an ungeared bypass fan model with solid metal rotor blades developed earlier by the authors are used for the mathematical model verification. Tip speed of rotor blades at the design point is Ucor. = 400m/s, bypass ratio — m = 8.4. Four booster stages are installed in the core duct. From first test results it is clear that required values of key parameters are achieved. Comparison of measured and calculated data gives evidence of their good agreement. At present, detailed tests of this fan and a similar fan with 3 booster stages are under way in the anechoic chamber of the CIAM’s C-3A acoustic test facility. The new direct-driven fan model described in this paper has quite different design values of parameters, geometry of the meridian contours, and shapes of outer and inner ducts. Tip speed of its rotor blades is reduced by 30 m/s, the hub diameter is decreased, and bypass ratio is higher (m = 11). In the near future, these two new models of non-geared and geared fans can be manufactured and tested.

Present paper contains application of inverse problem for 3D Navier-Stokes equations to design turbomachinery bladed rows. In-house software package used to solve 3D inverse problem is named 3D-INVERSE.EXBL. Inverse problem is based on desired static pressure distribution on suction side of blade, given blade thickness and pressure difference (named loading) in corresponding points of suction and pressure sides of blade. Inlet and outlet gas-dynamic parameters (pressure, density and flow velocity vector) are taken from direct solution of flow within multistage compressor and remains unchanged during inverse problem solution. Solution of inverse problem is determined using moving grid. 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. In the paper inverse solution provides effectiveness and operability of first rotor of multistage low pressure compressor (LPC) for a wide range of rpm (70 ± 100%) in case of absence of inlet guide vane (IGV).