Effect of Geometric Uncertainty on a One Stage Transonic Compressor of an Industrial Gas Turbine

2020 ◽  
Author(s):  
S. Venkatesh ◽  
K. Suzuki ◽  
M. Vahdati ◽  
L. Salles ◽  
Q. Rendu
Keyword(s):  
Leading Edge ◽  
Rotor Blades ◽  
Transonic Compressor ◽  
Manufacturing Tolerance ◽  
Geometric Uncertainty ◽  
Zigzag Pattern ◽  
Sinusoidal Pattern ◽  
And Performance ◽  
Edge Damage ◽  
Industrial Gas Turbine

Abstract The geometrical uncertainties can result in flow asymmetry around the annulus of compressor which in turn can detrimentally affect on the compressor stability and performance. Typically these uncertainties arise as a consequence of in-service degradation and/or manufacturing tolerance, both of which have been dealt with in this paper. The paper deals with effects of leading edge damage and tip gap on rotor blades. It was found that the chord-wise damage is more critical than radial damage. It was found that a zigzag pattern of arranging the damaged rotor blades (i.e. most damaged blades between two least damaged blades) would give the best possible performance and stability when performing maintenance and overhauling while a sinusoidal pattern of arrangement had the worst performance and stability. This behaviour of zigzag arrangement of random damaged blades is consonant with the behaviour of zigzag arrangement in random tip gaps. It is also shown in this work that the level of damage has a bigger impact on the compressor performance and stability than the number of damaged blades.

Effect of Geometric Uncertainty on a One Stage Transonic Compressor of an Industrial Gas Turbine

2021 ◽  
Author(s):  
Venkatesh Suriyanarayanan ◽  
Kentaro Suzuki ◽  
Mehdi Vahdati ◽  
Loic Salles ◽  
Quentin Rendu
Keyword(s):  
Gas Turbine ◽  
Transonic Compressor ◽  
One Stage ◽  
Geometric Uncertainty ◽  
Industrial Gas Turbine

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

2002 ◽  
Vol 124 (3) ◽  
pp. 351-357 ◽  
Author(s):  
William B. Roberts ◽  
Albert Armin ◽  
George Kassaseya ◽  
Kenneth L. Suder ◽  
Scott A. Thorp ◽  
...  
Keyword(s):  
Leading Edge ◽  
Aerodynamic Performance ◽  
Chord Length ◽  
Rotor Blades ◽  
Civil Aviation ◽  
Turbofan Engines ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Significant Difference ◽  
Leading Edges

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94–96% of nominal (new) blade chord.

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

2001 ◽  
Author(s):  
William B. Roberts ◽  
Albert Armin ◽  
George Kassaseya ◽  
Kenneth L. Suder ◽  
Scott A. Thorp ◽  
...  
Keyword(s):  
Leading Edge ◽  
Aerodynamic Performance ◽  
Chord Length ◽  
Rotor Blades ◽  
Civil Aviation ◽  
Turbofan Engines ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Significant Difference ◽  
Leading Edges

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100%, 80% and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

Numerical Modeling of the Effects of Leading-Edge Erosion and Trailing-Edge Damage on Wind Turbine Loads and Performance

2021 ◽  
Author(s):  
Francesco Papi ◽  
Lorenzo Cappugi ◽  
Sebastian Perez-Becker ◽  
Alessandro Bianchini
Keyword(s):  
Numerical Modeling ◽  
Wind Turbine ◽  
Leading Edge ◽  
Trailing Edge ◽  
And Performance ◽  
Edge Damage

On the Aerodynamic Design of Compressor Airfoils for Robustness Under Geometric Uncertainty

2004 ◽  
Author(s):  
Victor E. Garzon ◽  
David L. Darmofal
Keyword(s):  
Boundary Layer ◽  
Leading Edge ◽  
Suction Side ◽  
Loss Coefficient ◽  
Aerodynamic Design ◽  
Minimum Loss ◽  
Transonic Compressor ◽  
Geometric Uncertainty ◽  
The Mean ◽  
The Impact

This paper considers the aerodynamic design of compressor blade sections for improved performance robustness in the face of geometric uncertainty caused by noisy manufacturing processes. A probabilistic, gradient-based optimization method was used to redesign subsonic and transonic compressor airfoils subjected to geometric variability. Three different design goals were considered: Minimizing the deterministic profile total pressure loss coefficient, minimizing the mean value of loss coefficient, and minimizing the loss variability. In both transonic and subsonic applications, deterministic minimization of loss coefficient produced essentially the same airfoils as the probabilistic minimization of mean loss. However probabilistic minimization of loss variability produced clearly different airfoils which achieved reductions of 20% or more in standard deviation of loss compared to the minimum-loss designs. For the subsonic application, the improved robustness of the minimum variability design was achieved through a reduction in diffusion immediately downstream of the leading edge on the pressure side. This reduction in diffusion resulted in less sensitivity of the boundary layer to geometric variability in the leading edge region. For the transonic application, the robustness improvement was achieved by redesigning the suction side to produce a constant pressure region immediately downstream of the passage shock, which had the effect of desensitizing the boundary layer to variability in shock strength and position. A meanline model was used to assess the impact of probabilistic airfoil section optimization on overall compressor performance. While the mean efficiency was found to be nearly the same for all designs, the robust blade designs produced a decrease in compressor efficiency variability of 50% compared to the minimum-loss designs.

scholarly journals An Experimental Investigation of the Convective Heat Transfer on a Small Helicopter Rotor with Anti-Icing and De-Icing Test Setups

Aerospace ◽  
2021 ◽  
Vol 8 (4) ◽  
pp. 96
Author(s):  
Abdallah Samad ◽  
Eric Villeneuve ◽  
Caroline Blackburn ◽  
François Morency ◽  
Christophe Volat
Keyword(s):  
Heat Transfer ◽  
Convective Heat Transfer ◽  
Convective Heat ◽  
Leading Edge ◽  
Liquid Water Content ◽  
Rotor Blades ◽  
Direct Result ◽  
Average Error ◽  
Good Prediction ◽  
Series Of Experiments

Successful icing/de-icing simulations for rotorcraft require a good prediction of the convective heat transfer on the blade’s surface. Rotorcraft icing is an unwanted phenomenon that is known to cause flight cancelations, loss of rotor performance and severe vibrations that may have disastrous and deadly consequences. Following a series of experiments carried out at the Anti-icing Materials International Laboratory (AMIL), this paper provides heat transfer measurements on heated rotor blades, under both the anti-icing and de-icing modes in terms of the Nusselt Number (Nu). The objective is to develop correlations for the Nu in the presence of (1) an ice layer on the blades (NuIce) and (2) liquid water content (LWC) in the freestream with no ice (NuWet). For the sake of comparison, the NuWet and the NuIce are compared to heat transfer values in dry runs (NuDry). Measurements are reported on the nose of the blade-leading edge, for three rotor speeds (Ω) = 500, 900 and 1000 RPM; a pitch angle (θ) = 6°; and three different radial positions (r/R), r/R = 0.6, 0.75 and 0.95. The de-icing tests are performed twice, once for a glaze ice accretion and another time for rime ice. Results indicate that the NuDry and the NuWet directly increased with V∝, r/R or Ω, mainly due to an increase in the Reynolds number (Re). Measurements indicate that the NuWet to NuDry ratio was always larger than 1 as a direct result of the water spray addition. NuIce behavior was different and was largely affected by the ice thickness (tice) on the blade. However, the ice acted as insulation on the blade surface and the NuIce to NuDry ratio was always less than 1, thus minimizing the effect of convection. Four correlations are then proposed for the NuDry, the NuWet and the NuIce, with an average error between 3.61% and 12.41%. The NuDry correlation satisfies what is expected from heat transfer near the leading edge of an airfoil, where the NuDry correlates well with Re0.52.

scholarly journals Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

Energies ◽  
2021 ◽  
Vol 14 (14) ◽  
pp. 4168
Author(s):  
Botao Zhang ◽  
Xiaochen Mao ◽  
Xiaoxiong Wu ◽  
Bo Liu
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Axial Flow ◽  
Leading Edge ◽  
Rotating Stall ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
Transonic Compressor

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

Performance Prediction of Gas Turbine Under Different Strategies Using Low Heating Value Fuel

2013 ◽  
Author(s):  
Edson Batista da Silva ◽  
Marcelo Assato ◽  
Rosiane Cristina de Lima
Keyword(s):  
Natural Gas ◽  
Gas Turbine ◽  
Equivalence Ratio ◽  
Safe Operation ◽  
Engine Operation ◽  
Heating Value ◽  
Gasification Process ◽  
Surge Margin ◽  
And Performance ◽  
Industrial Gas Turbine

Usually, the turbogenerators are designed to fire a specific fuel, depending on the project of these engines may be allowed the operation with other kinds of fuel compositions. However, it is necessary a careful evaluation of the operational behavior and performance of them due to conversion, for example, from natural gas to different low heating value fuels. Thus, this work describes strategies used to simulate the performance of a single shaft industrial gas turbine designed to operate with natural gas when firing low heating value fuel, such as biomass fuel from gasification process or blast furnace gas (BFG). Air bled from the compressor and variable compressor geometry have been used as key strategies by this paper. Off-design performance simulations at a variety of ambient temperature conditions are described. It was observed the necessity for recovering the surge margin; both techniques showed good solutions to achieve the same level of safe operation in relation to the original engine. Finally, a flammability limit analysis in terms of the equivalence ratio was done. This analysis has the objective of verifying if the combustor will operate using the low heating value fuel. For the most engine operation cases investigated, the values were inside from minimum and maximum equivalence ratio range.

Influence of Leading Edge Fillet and Nonaxisymmetric Contoured Endwall on Turbine NGV Exit Flow Structure and Interactions With the Rim Seal Flow

2013 ◽  
Author(s):  
Özhan H. Turgut ◽  
Cengiz Camcı
Keyword(s):  
Flow Structure ◽  
Total Pressure ◽  
Guide Vane ◽  
Axial Flow ◽  
Leading Edge ◽  
Turbine Rotor ◽  
Rotor Blades ◽  
Constant Thickness ◽  
Computational Evaluation ◽  
Nozzle Guide

Three different ways are employed in the present paper to reduce the secondary flow related total pressure loss. These are nonaxisymmetric endwall contouring, leading edge (LE) fillet, and the combination of these two approaches. Experimental investigation and computational simulations are applied for the performance assessments. The experiments are carried out in the Axial Flow Turbine Research Facility (AFTRF) having a diameter of 91.66cm. The NGV exit flow structure was examined under the influence of a 29 bladed high pressure turbine rotor assembly operating at 1300 rpm. For the experimental measurement comparison, a reference Flat Insert endwall is installed in the nozzle guide vane (NGV) passage. It has a constant thickness with a cylindrical surface and is manufactured by a stereolithography (SLA) method. Four different LE fillets are designed, and they are attached to both cylindrical Flat Insert and the contoured endwall. Total pressure measurements are taken at rotor inlet plane with Kiel probe. The probe traversing is completed with one vane pitch and from 8% to 38% span. For one of the designs, area averaged loss is reduced by 15.06%. The simulation estimated this reduction as 7.11%. Computational evaluation is performed with the rotating domain and the rim seal flow between the NGV and the rotor blades. The most effective design reduced the mass averaged loss by 1.28% over the whole passage at the NGV exit.

Characterization of Shock Wave–Endwall Boundary Layer Interactions in a Transonic Compressor Rotor

1988 ◽  
Vol 110 (3) ◽  
pp. 386-392 ◽  
Author(s):  
D. C. Rabe ◽  
A. J. Wennerstrom ◽  
W. F. O’Brien
Keyword(s):  
Boundary Layer ◽  
Shock Wave ◽  
Air Force ◽  
Leading Edge ◽  
Transonic Compressor ◽  
Laser Measurements ◽  
Compressor Rotor ◽  
Boundary Layer Interaction ◽  
Air Force Base

The passage shock wave–endwall boundary layer interaction in a transonic compressor was investigated with a laser transit anemometer. The transonic compressor used in this investigation was developed by the General Electric Company under contract to the Air Force. The compressor testing was conducted in the Compressor Research Facility at Wright-Patterson Air Force Base, OH. Laser measurements were made in two blade passages at seven axial locations from 10 percent of the axial blade chord in front of the leading edge to 30 percent of the axial blade chord into the blade passage. At three of these axial locations, laser traverses were taken at different radial immersions. A total of 27 different locations were traversed circumferentially. The measurements reveal that the endwall boundary layer in this region is separated from the core flow by what appears to be a shear layer where the passage shock wave and all ordered flow seem to end abruptly.

Sign in / Sign up

Export Citation Format

Share Document