Heat Transfer Analysis for Servo Valve in Hydraulic Servo Actuator of Aero Engine Vectoring Nozzle

2016 ◽  
Author(s):  
Yulin Ding ◽  
Youhong Liu ◽  
Yifu Luo ◽  
Liwei Du
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Analysis ◽  
Servo Valve ◽  
Aero Engine ◽  
Servo Actuator ◽  
Hydraulic Servo

Design and Testing of a Rig to Investigate Buoyancy-Induced Heat Transfer in Aero-Engine Compressor Rotors

2020 ◽  
Author(s):  
Dario Luberti ◽  
Marios Patinios ◽  
Richard W. Jackson ◽  
Hui Tang ◽  
Oliver J. Pountney ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Radial Distribution ◽  
Tip Clearance ◽  
Heat Transfer Analysis ◽  
Radiant Heat Transfer ◽  
Central Cavity ◽  
Engine Operation ◽  
Closed Cavity ◽  
Rotating Discs ◽  
Aero Engine

Abstract The change in compressor blade-tip clearance across the flight cycle depends on the expansion of the rotor, which in turn depends on the temperature and stress in the discs. The radial distribution of temperature is directly coupled to the buoyancy-driven flow and heat transfer in the rotating disc cavities. This paper describes a new test rig specifically designed to investigate this conjugate phenomenon. The rig test section includes four rotating discs enclosing three cavities. Two discs in the central cavity are instrumented with thermocouples to provide the radial distribution of temperature; the two outer cavities are thermally insulated to create appropriate boundary conditions for the heat transfer analysis. An axial throughflow of air is supplied between a stationary shaft and the bore of the discs. The temperature of the throughflow air is measured by thermocouples in rakes upstream and downstream of the central cavity. For a cold throughflow, the outer shroud of the central cavity is heated. Two independently-controlled radiant heaters allow differential shroud temperatures for the upstream and downstream discs, as found in aero-engine compressors. Alternatively, the throughflow can be heated above the shroud temperature to simulate the transient conditions during engine operation where stratified flow can occur inside the cavity. The rig is designed to operate in conditions where both convective and radiative heat transfer dominate; all internal surfaces of the cavity are painted matt black to allow the accurate calculation of the radiant heat transfer. Separate attachments can be fitted to the cobs of both central discs; the attachments reduce the axial gap between the cobs — reducing the gap to zero creates a closed cavity, which can occur in some compressor designs. Other instrumentation includes heat-flux gauges on the shroud and high-frequency pressure transducers embedded into the disc diaphragm to capture unsteady flow structures. Attention has been given to experimental uncertainty, including the computation of the thermal-disturbance errors, caused by thermocouples embedded in the rotating discs; a Bayesian statistical model is used to reduce the effect of uncertainties in temperature measurements on the calculation of the Nusselt number. The effect of relevant non-dimensional parameters on the radial distribution of the disc and throughflow temperatures has been shown for some typical cases.

Design and Testing of a Rig to Investigate Buoyancy-Induced Heat Transfer in Aero-Engine Compressor Rotors

2020 ◽  
Author(s):  
Dario Luberti ◽  
Marios Patinios ◽  
Richard Jackson ◽  
Hui Tang ◽  
Oliver J Pountney ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Radial Distribution ◽  
Radiant Heat ◽  
Tip Clearance ◽  
Heat Transfer Analysis ◽  
Radiant Heat Transfer ◽  
Rotating Disc ◽  
Central Cavity ◽  
Engine Operation ◽  
Aero Engine

Abstract The change in compressor blade-tip clearance across the flight cycle depends on the expansion of the rotor, which in turn depends on the temperature and stress in the discs. The radial distribution of temperature is directly coupled to the buoyancy-driven flow and heat transfer in the rotating disc cavities. This paper describes a new test rig specifically designed to investigate this conjugate phenomenon. The rig test section includes four rotating discs enclosing three cavities. Two discs in the central cavity are instrumented with thermocouples to provide the radial distribution of temperature; the two outer cavities are thermally insulated to create appropriate boundary conditions for the heat transfer analysis. An axial throughflow of air is supplied between a stationary shaft and the bore of the discs. The temperature of the throughflow air is measured by thermocouples in rakes upstream and downstream of the central cavity. For a cold throughflow, the outer shroud of the central cavity is heated. Two independently-controlled radiant heaters allow differential shroud temperatures for the upstream and downstream discs, as found in aero-engine compressors. Alternatively, the throughflow can be heated above the shroud temperature to simulate the transient conditions during engine operation where stratified flow can occur inside the cavity. The rig is designed to operate in conditions where both convective and radiative heat transfer dominate; all internal surfaces of the cavity are painted matt black to allow the accurate calculation of the radiant heat transfer.

Conjugate Heat Transfer Analysis of Military Aero Engine Combustor Liner With Impingement and Effusion Cooling

2019 ◽  
Author(s):  
Batchu Suresh ◽  
Chinmayee Panigrahi ◽  
Antonio Davis ◽  
V. Kesavan ◽  
D. Kishore Prasad
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Hole Diameter ◽  
Heat Transfer Analysis ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Impingement Plate ◽  
Aero Engine ◽  
Combustor Liner

Abstract Improvement in specific thrust is one of the desirable requirements for Military aero-engine which has led to a tremendous increase in turbine inlet temperatures. This has resulted in combustion chambers to operate at a gas temperature of the order of 2100K, making it difficult for the thermal designers to design a liner cooling configuration to bring down its metal temperature within allowable limits with available coolant air. The present study highlights the computational prediction of cooling effectiveness for impingement-effusion cooled combustor liner. The impingement cooling is adopted to the effusion cooled liner in order to enhance its coolant side heat transfer. 1-Dimensional (1-D) analysis is carried out to obtain a suitable impingement geometry to improve the coolant side heat transfer. Suitable geometrical features like impingement hole diameter (di) and distance of the impingement plate from effusion liner (z) are arrived for enhancement of coolant side heat transfer. Conjugate Heat Transfer analysis (CHT) is carried out for three cooling configurations with different impingement hole diameter. Effusion cooled liner with porosity 1% and holes inclined at 22° and for impingement plate hole porosity of 1.6% is maintained for all the configurations. CHT analysis is carried out for effusion cooled liner using ANSYS Fluent ver.14.5. The film cooling predictions are in good agreement for effusion cooled liner plate with measurements. SST k-ω turbulence model with enhanced wall function predicted well. The effectiveness obtained for effusion cooled liner and impingement-effusion cooled liner are compared. There is an improvement of 34% in effectiveness for impingement-effusion cooled liner compared to effusion cooled liner with a reduction of coolant air mass flow by 10%. The variation of temperature for the impingement-effusion cooled liner is lower. Parametric analysis is also carried out to study the effect of blowing ratio and metal thermal conductivity on the film cooling effectiveness for impingement-effusion cooled liner.

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