Experimental and Numerical Studies on Highly Loaded Supersonic Axial Turbine Cascades

2010 ◽  
Author(s):  
T. Wolf ◽  
F. Kost ◽  
E. Janke ◽  
F. Haselbach ◽  
L. Willer
Keyword(s):  
High Pressure ◽  
Mach Number ◽  
Attractive Alternative ◽  
High Pressure Turbine ◽  
Number Range ◽  
Potential Benefits ◽  
High Mach ◽  
Aero Engines ◽  
Turbine Airfoils ◽  
Highly Loaded

For the small to medium thrust range of modern aero engines, highly loaded single stage HP turbines facilitate an attractive alternative to a more conventional 2-stage HPT architecture. Whereas the potential benefits of reductions in component length and part count, hence, in weight and cost do motivate their application, the related risks are in maintaining associated losses of supersonic flows at low values as well as managing the interaction losses between HPT and the downstream sub-component to arrive at competitive levels of component efficiencies. This paper focuses on fundamental aerodynamic concept studies and related cascade experiments in support of a future highly loaded high-pressure turbine architecture. Starting with some general remarks on low-loss supersonic aerodynamic concepts for high-pressure turbines, results from development efforts towards 2D airfoil concepts viable for high-pressure turbine airfoils are shown. In particular, CFD based design approaches are compared against experimental data taken at DLR Go¨ttingen in un-cooled cascade tests and at engine representative levels of Mach and Reynolds numbers. For the airfoils investigated, it turns out that there is indeed a supersonic Mach number range were loss levels are comparable to high Mach number subsonic values, thereby enabling a competitive aerodynamic design concept for a 3D high-pressure turbine stage.

Design and Test of a Highly Loaded Single-Stage High Pressure Turbine

2011 ◽  
Author(s):  
T. Wolf ◽  
E. Janke ◽  
R. Benton ◽  
F. Kost ◽  
F. Haselbach ◽  
...  
Keyword(s):  
High Pressure ◽  
Single Stage ◽  
Medium Size ◽  
Attractive Alternative ◽  
Efficient Design ◽  
High Pressure Turbine ◽  
Comparison Of Experiments ◽  
Performance Area ◽  
Aero Engines ◽  
Highly Loaded

For the small to medium thrust range of modern aero engines, highly loaded single-stage HP turbines facilitate an attractive alternative to a more conventional 2-stage HPT architecture. Within the German government funded LUFO-3 programme “Transonic Single Stage High-Pressure Turbine”, a substantial activity towards the development and test of supersonic aerodynamic technology for single stage turbines was launched in 2003. This paper describes fundamental aerodynamic concept studies and related cascade experiments in support of a future highly loaded high-pressure turbine architecture. Details of the first out of two builds featuring an engine representative single-stage HPT is described in detail. Focus will be on instrumentation design, selected results from performance, area traverse and unsteady blade surface pressure measurements and the comparison of experiments with numerical simulations. The successfully completed test campaign confirms the existence of an aerodynamically efficient design of a highly loaded HPT, thereby enabling a competitive building block for a small to medium size engine concept.

Investigation of Trailing Edge Cooling Concepts in a High Pressure Turbine Cascade—Aerodynamic Experiments and Loss Analysis

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Hans-Ju¨rgen Rehder
Keyword(s):  
Boundary Layer ◽  
High Pressure ◽  
Coolant Flow ◽  
Coolant Fluid ◽  
Pressure Side ◽  
Turbine Cascade ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Number Range ◽  
Mixing Losses

As part of a European research project, the aerodynamic and thermodynamic performance of a high pressure turbine cascade with different trailing edge cooling configurations was investigated in the wind tunnel for linear cascades at DLR in Go¨ttingen. A transonic rotor profile with a relative thick trailing edge was chosen for the experiments. Three trailing edge cooling configurations were applied, first central trailing edge ejection, second a trailing edge shape with a pressure side cut-back and slot equipped with a diffuser rib array, and third pressure side film cooling through a row of cylindrical holes. For comparison, aerodynamic investigations on a reference cascade with solid blades (no cooling holes or slots) were performed. The experiments covered the subsonic, transonic and supersonic exit Mach number range of the cascade while varying cooling mass flow ratios up to 2 %. This paper analyzes the effect of coolant ejection on the airfoil losses. Emphasis was given on separating the different loss contributions due to shocks, pressure, and suction side boundary layer, trailing edge, and mixing of the coolant flow. Employed measurement techniques are schlieren visualization, blade surface pressure measurements, and traverses by pneumatic probes in the cascade exit flow field and around the trailing edge. The results show that central trailing edge ejection significantly reduces the mixing losses and therefore decreases the overall loss. Higher loss levels are obtained when applying the configurations with pressure side blowing. In particular, the cut-back geometry reveals strong mixing losses due to the low momentum coolant fluid, which is decelerated by the diffuser rib array inside the slot. The influence of coolant flow rate on the trailing edge loss is tremendous, too. Shock and boundary layer losses are major contributions to the overall loss but are less affected by the coolant. Finally a parameter variation changing the temperature ratio of coolant to main flow was performed, resulting in increasing losses with decreasing coolant temperature.

Weibull Distributions Applied to Cost and Risk Analysis for Aero Engines

2008 ◽  
Author(s):  
D. S. Pascovici ◽  
K. G. Kyprianidis ◽  
F. Colmenares ◽  
S. O. T. Ogaji ◽  
P. Pilidis
Keyword(s):  
High Pressure ◽  
Risk Analysis ◽  
Economic Model ◽  
Spare Parts ◽  
Weibull Distributions ◽  
High Pressure Turbine ◽  
Aero Engines ◽  
The Cost ◽  
The Impact ◽  
Different Parts

This paper presents the use of Weibull formulation to the life analysis of different parts of the engine in order to estimate the cost of maintenance, the direct operating costs (DOC) and net present cost (NPC) of future type turbofan engines. The Weibull distribution is often used in the field of life data analysis due to its flexibility—it can mimic the behavior of other statistical distributions such as the normal and the exponential. The developed economic model is composed of three modules: a lifing module, an economic module and a risk module. The lifing module estimates the life of the high pressure turbine blades through the analysis of creep and fatigue over a full working cycle of the engine. The value of life calculated by the lifing is then taken as the baseline distribution to calculate the life of other important modules of the engine using the Weibull approach. Then the lower of the values of life of all the distributions is taken as time between overhaul (TBO), and used into the economic module calculations. The economic module uses the TBO together with the cost of labour and the cost of the engine (needed to determine the cost of spare parts) to estimate the cost of maintenance and DOC of the engine. In the present work five Weibull distributions are used for five important sources of interruption of the working life of the engine: Combustor, Life Limited Parts (LLP), High Pressure Compressor (HPC), General breakdowns and High Pressure Turbine (HPT). The risk analysis done in this work shows the impact of the breakdown of different parts of the engine on the NPC and DOC, the importance that each module of the engine has in its life, and how the application of the Weibull theory can help us in the risk assessment of future aero engines. A detailed explanation of the economic model is done in two other works (Pascovici et. al. [6] and Pascovici et. al. [7]), so in this paper only a general overview is done.

Delayed Detached-Eddy Simulations of a High Pressure Turbine Vane

2016 ◽  
Author(s):  
Dun Lin ◽  
Xinrong Su ◽  
Xin Yuan
Keyword(s):  
Reynolds Number ◽  
High Pressure ◽  
Mach Number ◽  
Shock Waves ◽  
The Other ◽  
Pressure Waves ◽  
High Pressure Turbine ◽  
Isothermal Boundary ◽  
Turbine Vane ◽  
Exit Mach Number

The flow in a generic, high-pressure turbine vane was simulated using an in-house DDES code. Two different operating conditions were simulated with one leading to a shock wave while the other does not. One case was used to validate the capability of the DDES method to capture shock waves and other flow structures using an inlet Reynolds number of 271,000 and an exit Mach number of 0.840. The test conditions for the other case were an inlet Reynolds number of 265,000 and an exit Mach number of 0.927, which is representative of a transonic, high pressure turbine vane which was used to further investigate the flow field. The DDES simulations from the first case are compared with published experimental data, RANS simulations and LES simulations. Then, DDES results for two cases with adiabatic and isothermal boundary conditions are compared. The numerical simulations with the isothermal boundary condition are further used to study the flow phenomena with wake vortices, shock waves, pressure waves, wake-shock interactions, and wake-pressure wave interactions. The effects of the pressure waves on the vane heat transfer are also analyzed.

An Investigation of Reynolds Lapse Rate for Highly Loaded Low Pressure Turbine Airfoils With Forward and Aft Loading

2011 ◽  
Author(s):  
M. Eric Lyall ◽  
Paul I. King ◽  
Rolf Sondergaard ◽  
John P. Clark ◽  
Mark W. McQuilling
Keyword(s):  
Reynolds Number ◽  
Eddy Viscosity ◽  
Low Pressure ◽  
Lapse Rate ◽  
Freestream Turbulence ◽  
Number Range ◽  
Low Pressure Turbine ◽  
Turbine Airfoils ◽  
Hot Film ◽  
Highly Loaded

This paper presents an experimental and computational study of the midspan low Reynolds number loss behavior for two highly loaded low pressure turbine airfoils, designated L2F and L2A, which are forward and aft loaded, respectively. Both airfoils were designed with incompressible Zweifel loading coefficients of 1.59. Computational predictions are provided using two codes, Fluent (with k-k1-ω model) and AFRL’s Turbine Design and Analysis System (TDAAS), each with a different eddy-viscosity RANS based turbulence model with transition capability. Experiments were conducted in a low speed wind tunnel to provide transition models for computational comparisons. The Reynolds number range based on axial chord and inlet velocity was 20,000 < Re < 100,000 with an inlet turbulence intensity of 3.1%. Predictions using TDAAS agreed well with the measured Reynolds lapse rate. Computations using Fluent however, predicted stall to occur at significantly higher Reynolds numbers as compared to experiment. Based on triple sensor hot-film measurements, Fluent’s premature stall behavior is likely the result of the eddy-viscosity hypothesis inadequately capturing anisotropic freestream turbulence effects. Furthermore, rapid distortion theory is considered as a possible analytical tool for studying freestream turbulence that influences transition near the suction surface of LPT airfoils. Comparisons with triple sensor hot-film measurements indicate that the technique is promising but more research is required to confirm its utility.

An Investigation of Reynolds Lapse Rate for Highly Loaded Low Pressure Turbine Airfoils With Forward and Aft Loading

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
M. Eric Lyall ◽  
Paul I. King ◽  
Rolf Sondergaard ◽  
John P. Clark ◽  
Mark W. McQuilling
Keyword(s):  
Reynolds Number ◽  
Eddy Viscosity ◽  
Low Pressure ◽  
Lapse Rate ◽  
Freestream Turbulence ◽  
Number Range ◽  
Low Pressure Turbine ◽  
Turbine Airfoils ◽  
Hot Film ◽  
Highly Loaded

This paper presents an experimental and computational study of the midspan low Reynolds number loss behavior for two highly loaded low pressure turbine airfoils, designated L2F and L2A, which are forward and aft loaded, respectively. Both airfoils were designed with incompressible Zweifel loading coefficients of 1.59. Computational predictions are provided using two codes, Fluent (with k-kl-ω model) and AFRL’s Turbine Design and Analysis System (TDAAS), each with a different eddy-viscosity RANS based turbulence model with transition capability. Experiments were conducted in a low speed wind tunnel to provide transition models for computational comparisons. The Reynolds number range based on axial chord and inlet velocity was 20,000 < Re < 100,000 with an inlet turbulence intensity of 3.1%. Predictions using TDAAS agreed well with the measured Reynolds lapse rate. Computations using Fluent however, predicted stall to occur at significantly higher Reynolds numbers as compared to experiment. Based on triple sensor hot-film measurements, Fluent’s premature stall behavior is likely the result of the eddy-viscosity hypothesis inadequately capturing anisotropic freestream turbulence effects. Furthermore, rapid distortion theory is considered as a possible analytical tool for studying freestream turbulence that influences transition near the suction surface of LPT airfoils. Comparisons with triple sensor hot-film measurements indicate that the technique is promising but more research is required to confirm its utility.

Study of Secondary Flows in Blade Cascades of Turbomachines

1982 ◽  
Vol 104 (2) ◽  
pp. 497-509 ◽  
Author(s):  
F. Bario ◽  
F. Leboeuf ◽  
K. D. Papailiou
Keyword(s):  
Mach Number ◽  
Turbine Blades ◽  
Theoretical Method ◽  
Secondary Flows ◽  
Compressor Blades ◽  
Number Range ◽  
Mean Values ◽  
Method Of Calculation ◽  
Passage Vortex ◽  
Highly Loaded

Experiments have been performed with two cascades of turbomachines. The first cascade is composed of highly loaded turbine blades, and has been used in the low subsonic Mach number range. The second cascade consists of highly loaded compressor blades, of the DCA type. The Mach number was then in the high subsonic range. The experimental results are presented in the form of mean values in the pitch direction. Detailed local values are also described. The growth of a passage vortex and a corner effect are presented in the compressor case. Their interactions with the whole flow are analyzed. In the turbine case, the passage vortex is found to be a dominant effect. Results obtained with a theoretical method of calculation of the flow in the blade passage are used to complete the analysis.

ON THE RESPONSE OF SHOCK-INDUCED SEPARATION BUBBLE TO SMALL AMPLITUDE DISTURBANCES

2005 ◽  
Vol 19 (28n29) ◽  
pp. 1495-1498 ◽  
Author(s):  
L. KRISHNAN ◽  
Y. YAO ◽  
N. D. SANDHAM ◽  
G. T. ROBERTS
Keyword(s):  
Mach Number ◽  
Small Amplitude ◽  
Separation Bubble ◽  
Constant Wall Temperature ◽  
Number Range ◽  
Upstream Side ◽  
Bubble Length ◽  
Streamwise Structures ◽  
High Mach ◽  
Recovery Temperature

Numerical simulations of an oblique shock interacting with a compressible laminar boundary layer are reported. The Mach number ranges from 2 to 6.85, while the Reynolds number based on the distance to the impingement location is fixed at 3 × 105. All the simulations are carried out with a constant wall temperature, equal to the adiabatic recovery temperature. At higher shock strength the evolved separation bubbles are taller and are biased towards the upstream side of the impingement location with an asymmetrical structure. Existing similarity scalings for the bubble length need to be modified for the high Mach number range. Introduction of small amplitude disturbances upstream of the separation bubble resulted in the growth of organised streamwise structures downstream of the bubble.

Unsteady Film Cooling Performance on the High Pressure Turbine Shroud Under Rotor-Stator Interaction for an Aero-Engine

2021 ◽  
Author(s):  
Zihao Bao ◽  
Zhihai Kou ◽  
Bo Han ◽  
Guangchao Li
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Leakage Flow ◽  
High Pressure Turbine ◽  
Cooling Performance ◽  
Blade Rotation ◽  
Aero Engine ◽  
Rotor Stator Interaction ◽  
Blade Tip ◽  
Aero Engines

Abstract The turbine rotor inlet temperature of modern aero-engines has continuously increased in order to achieve higher thrust-to-weight ratio and thermal efficiency, which requires higher cooling effectiveness for turbine components. The turbine shroud is exposed to the blade tip leakage flow, and has become a common limiting factor for the turbine stage of advanced aero-engines. The three-dimensional numerical simulation on the unsteady film cooling characteristics of the high-pressure turbine shroud for an aero-engine under the rotor-stator interaction and the high blade rotation speed was conducted. The sliding grid technology was used to realize the relative movement between the turbine blade and the turbine shroud, and the rotor-stator interaction. Effects of the blade rotation, blowing ratios and the film jet direction on the unsteady film cooling performance of the high-pressure turbine shroud were revealed. It is found that the film cooling characteristics of the turbine shroud present an unsteady and periodic phenomenon. The blade tip clearance leakage flow, leakage vortex and mainstream suppression have important effects on the film cooling performance of the high-pressure turbine shroud. More attention should be paid to the insufficient cooling margin of the front row film holes due to coolant jet liftoff from the shroud surface under the high blowing ratio.

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