Aerodynamics of Contra-Rotating Fans With Swept Blades

A Counter-Rotating System (CRS) is composed of a front rotor and a rear rotor which rotates in the opposite direction. Compared with traditional rotor-stator system, the rear rotor is used not only to recover the static head but also to supply energy to the fluid. Therefore, to achieve the same performance, the use of a CRS may lead to a reduction of the rotational speed and may generate better homogeneous flow downstream of the stage. On the other hand, the mixing area in between the two rotors induces complicated interacting flow structures. Blade sweep has attracted the turbomachinery blade designers owing to a variety of performance benefits it offers. However, the effect of blade sweep on the performance, stall margin improvements whether it is advantageous/disadvantageous to sweep one or both rotors has not been studied till now. In the current investigation blade sweep on the performance characteristics of contra rotating axial flow fans are studied. Two sweep schemes (axial sweeping and tip chord line sweeping) are studied for two sweep angles (20° and 30°). Effect of blade sweep on front rotor and rear rotor are dealt separately by sweeping one at a time. Both rotors are swept together and effect of such sweep scheme on the aerodynamic performance of the stage is also reported here. The performance of contra rotating fan is significantly affected by all these parameters. Blade sweep improved the pressure rise and stall margin of front rotors. Axially swept rotors are found to have higher pressure rise with reduced incidence losses near the tip for front rotors. Sweeping the rear rotor is not effective since the pressure rise is less than that of unswept rotor and also has less stall margin.

Studies on Characteristic Frequency and Length Scale of Shock Induced Motion in Transonic Diffuser Using a High Order LES Approach

Unsteady transonic flows in diffuser have become increasingly important, because of its application in new propulsion systems. In the development of supersonic inlet, air breathing propulsion systems of aircraft and missiles, detail investigations of these types of flow behavior are very much essential. In these propulsion systems, naturally present self-sustaining oscillations, believed to be equivalent to dynamically distorted flow fields in operational inlets, were found under all operating conditions. The investigations are also relevant to pressure oscillations known to occur in ramjet inlets in response to combustor instabilities. The unsteady aspects of these flows are important because the appearance of undesirable fluctuations generally impose limitation on the inlet performance. Test results of ramjet propulsion systems have shown undesirable high amplitude pressure fluctuations caused by the combustion instability. The pressure fluctuations originated from the combustor extend forward into the inlet and interact with the diffuser flow-field. Depending on different parameters such as the diffuser geometry, the inlet/exit pressure ratio, the flow Mach number, different complicated phenomena may occur. The most important characteristics are the occurrence of shock induced separation, the length of separation region downstream of the shock location, and the oscillation of shock location as well as the oscillation of the whole downstream flow. Sajben experimentally investigated in detail the time mean and unsteady flow characteristics of supercritical transonic diffuser as a function of flow Mach number upstream the shock location and diffuser length. The flows exhibited features similar to those in supersonic inlets of air-breathing propulsion systems of aircraft. A High-order LES turbulence model developed by the author is assessed with experimental data of Sajben on the self-excited shock oscillation phenomena. The whole diffuser model configuration including the suction slot located at certain axial location around the bottom and side walls to remove boundary layer, are included in the present computation model. The time-mean and unsteady flow characteristics in this transonic diffuser as a function of flow Mach number and diffuser length are investigated in detail. The results of study showed that in the case of shock-induced separation flow, the length and thickness of the reverse flow region of the separation-bubble change, as the shock passed through its cycle. The instabilities in the separated layer, the shock /boundary layer interaction, the dynamics of entrainment in the separation bubble, and the interaction of the travelling pressure wave with the pressure fluctuation region caused by the step-like structure of the suction slot play very important role in the shock-oscillation frequency.

Assessment of Different Turbulence Models for Predicting Laminar Separation Bubble Over Thick Airfoils

Accurate laminar-turbulent prediction is very much important to understand the complete performance characteristics of any airfoil which operates at low and medium Reynolds number. In this article, a numerical study has been performed over two different thick airfoils operating at low Reynolds number using k-ω SST, k-kl-ω and Spalart-Allmaras (SA) RANS models. The unsteady two dimensional (2D) simulations are performed over NACA 0021 and NACA 65-021 at Re 120,000 for a range of angle of attacks. The performances of these models are assessed through aerodynamic lift, drag and pressure coefficients. To obtain better comparison, the simulated results are compared with the experimental measurements and XFOIL results as well. In this present study, it is found that the k-kl-ω transition model is capable of predicting correct lift, drag coefficient and separation bubble as reported in experiments. At high angles of attack, this model fails to predict performance variables accurately. The SA and SST models are fail to predict laminar separation bubble. However, At high angle of attack, SA model shows better predictions compared to k-kl-ω and k-ω SST models.

Non-Linear Transient Stability Analysis of a Rigid Rotor Supported on Journal Bearings With Rectangular Dimples

The influence of deterministic surface texture on the sub-synchronous whirl stability of a rigid rotor has been studied. Non-linear transient stability analysis has been performed to study the stability of a rigid rotor supported on two symmetric journal bearings with a rectangular dimple of large aspect ratio. The surface texture parameters considered are dimple depth to minimum film thickness ratio and the location of the dimple on the bearing surface. Journal bearings of different Length to diameter ratios have been studied. The governing Reynolds equation for finite journal bearings with incompressible fluid has been solved using the Finite Element Method under isothermal conditions. The trajectories of the journal center have been obtained by solving the equations of motion of the journal center by the fourth-order Runge-Kutta method. When the dimple is located in the raising part of the pressure curve the positive rectangular dimple is seen to decrease the stability whereas the negative rectangular dimple is seen to improve the stability of the rigid rotor.

Numerical Analysis of Impingement/Effusion Cooling Effectiveness on Flat Plates

The impingement/effusion cooling is a method of using cooling air to protect the hot combustor liner surfaces from high temperature effectively. This paper investigates the impingement/effusion cooling over two perforated flat plates and proposes a better cooling scheme for high temperature combustion liners in aircrafts and electrical power generation application. The adiabatic cooling effectiveness distribution over the liner surface is numerically studied by control volume technique in CFD. In this hybrid scheme the hydraulic diameter (d) of the hole is 1mm and impingement plate is provided with holes normal to the plate over its whole length of 250d. While effusion plate has only 20 rows of holes inclined at 30° to its surface. The effect of blowing ratio (BR) over this hybrid scheme of cooling is studied for different BR of 0.5, 1.0, 1.5 and 2.0. It has been found that the area averaged effectiveness increases steeply for BR 0.5 to 1.0 but further increase in BR results only in a small increase. The results also show that increasing the hole diameter increases averaged effectiveness while increasing the center-to-center spacing decreases averaged effectiveness.

CFD Analysis to Investigate the Effect of Vortex Generators on a Transonic Axial Flow Compressor Stage

The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. Computational Fluid Dynamics (CFD) has been extensively used to analyze the flow through rotating machineries, in general and through axial compressors, in particular. The present work is focused on the studying the effects of Vortex Generator (VG) on test compressor at CSIR National Aerospace Laboratories, Bangalore, India using CFD. The compressor consists of NACA transonic rotor with 21 blades and subsonic stator with 18 vanes. The design pressure ratio is 1.35 at 12930 RPM with a mass flow rate of 22 kg/s. Three configurations of counter rotating VGs were selected for the analysis with 0.25δ, 0.5δ and δ height, where δ was equal to the physical thickness of boundary layer (8mm) at inlet to the compressor rotor [11]. The vortex generators were placed inside the casing at 18 percent of the chord ahead to the leading edge of the rotor. A total of 63 pairs of VGs were incorporated, with three pairs in one blade passage. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 3.5% from baseline at design speed. The reasons for this numerical stall margin improvement are discussed in detail.

Multiaxial Fatigue Life Estimation in the Absence of Fatigue Properties: A Case Study on a Turbine Rotor Used in a Typical Turbo Shaft Engine

Complex stress strain response of a turbine rotor used in a gas turbine engine was studied. Simple and comprehensive approximation techniques developed by Muralidharan–Manson, Bäumel-Seeger (from data obtained from tension tests) and Roessle–Fatemi (from data obtained from hardness tests) were used to predict the fatigue constants of the rotor material. Multiaxial Fatigue damage models like von Mises equivalent strain model, Smith Watson Topper model, Fatemi–Socie Model, Kandil Brown and Miller model were used to predict the fatigue life of the rotor. Predictions were then compared with the life obtained from the same damage models using the experimental fatigue constants and the life obtained from Low Cycle Fatigue (LCF) testing of the turbine rotor. Acceptable life predictions were obtained with SWT model and FS model using the fatigue constants obtained from the experiment as well as from the approximation techniques. von-Mises equivalent strain model failed to give reasonable life predictions with fatigue constants obtained from the experiment and approximation techniques. The life predicted by KBM model using fatigue constants obtained from approximation techniques (Bäumel-Seeger and Roessle-Fatemi) was found unsatisfactory. The approximation technique proposed by Muralidharan-Manson in combination with all the damage models fitted the failure data within a factor of 5. Finite Element tools were used to determine the stress/strain response of the component under the mutiaxial loading condition.

Shape Optimization of Flexible Supports for Aero Engines

Flexible supports are used in many aero and automobile industrial applications. They transmit loads, accommodate misalignment, allow axial displacement, ensure no loss of lubricants, absorb shock and dampen vibration, withstand high temperatures, allow easy installation and disassembling. Flexible supports react on connected equipment components when subjected to misalignment and torque. The reaction forces and moments on components due to flexible supports should be within the allowable limits or otherwise it can cause failure of gears, shafts, bearings, and other equipment components. These flexible supports used in aero engine applications expected to meet design and manufacturing criteria. Flexible supports should have required stiffness values in different directions to meet rotor dynamic stability criteria. Flexible supports also required to meet strength and durability criteria for the given material at the required maximum operating temperature. The designed component should be producible and meet manufacturing limitations. The main objective of this paper is to optimize single and multiple convolutes types of flexible supports with in the manufacturing limits and in the given design space. A methodology is developed to optimize the components to meet required stiffness, strength and durability criteria. Parametric models of flexible support are developed in UNIGRAPHICS NX9. Design parameters such as overall length, convolute height, convolute radius and angle are considered for the optimization study. ANSYS Workbench is used for the analysis and optimization of flexible support.

Numerical Computation of a Turbulent Lifted Flame Using Flamelet Generated Manifold With Different Progress Variable Definitions

Dimension reduction is a popular and attractive approach for modeling turbulent reacting flow incorporating finite rate chemistry effects. One of the earliest and most popular approaches in this category is the Laminar Flamelet Model (LFM), which represents the turbulent flame brush using statistical averaging of laminar flamelets whose structure is not affected by turbulence. The other common reduction approach is the intrinsic low dimensional manifold (ILDM). While, the LFM has limitations in predicting the non-equilibrium effects, the ILDM model suffers in the prediction of the low temperature kinetics. A combination of the two approaches where flamelet based manifold are generated called, Flamelet Generated Manifold (FGM) model considers that the scalar evolution in a turbulent flame can be approximated by the scalar evolution similar to that in a laminar flame. This model does not involve any assumption on flame structure. Therefore, it can be successfully used to model ignition, slow chemistry and quenching effects, which are far away from equilibrium. In the FGM, the manifold can be created using different flame configurations. For premixed flames, 1D unstrained flamelets are solved in reaction-progress space. In the case of diffusion flames, a counter flow configuration is used to generate a series of steady flamelets with increasing scalar dissipation and also an unsteady laminar flamelet is generated to create the diffusion FGM manifold. In the present work, a diffusion flamelet based FGM model is compared with the FGM model using premixed unstrained flamelet configurations. The performance and predictive capabilities of the two approaches are compared for a turbulent lifted methane flame in a diluted hot co-flow environment, where the reacting flow associated with the central jet exhibits similar chemical kinetics, heat transfer and molecular transport as recirculation burners without the complex recirculating fluid structures. It is observed that though the diffusion flamelet based FGM predicts a lifted flame, but the lift off height is lower compared to the premixed configuration. A parametric study with different normalization for the progress variable is done to study its impact on the flame characteristics and the manifold created. Finally, the computations are performed for different definitions of the progress variable from previously published works. It is seen that the results are sensitive to the various progress variable definitions, particularly when the number of species are higher and involve different time scales.

Validation of a 1D Transient Simulation Model of a Multistage Axial Compressor

An unsteady one-dimensional dynamic model has been developed at Georgia Tech to investigate the impact of stage characteristics as well as load distribution on the compression and expansion waves that develop prior to a surge event in a multistage axial compressor. In the developed model, each of the blade rows is replaced by a duct of varying cross-sectional area with force and work source terms. The source terms model the force and energy imparted by a blade row to the working fluid. The modeling assumes the flow to be inviscid, unsteady, compressible and axisymmetric. While rotating stall cannot be explicitly modeled in a 1D mean-line method, the effect of rotating stall can be captured by a judicious choice of source terms that reflects the loss of pumping capability of a stage. Conservation of mass, momentum and energy are applied to an elemental control volume resulting in one-dimensional quasi-linear Euler system of equations. A non-uniform grid and the second-order central difference Kurganov-Tadmor (KT) scheme are used to discretize the one-dimensional computational domain. The resulting ODEs are solved with an explicit second order Runge-Kutta solver. A throttle schedule is used to introduce perturbations at a selected operating condition in order to study flow oscillations that can lead to a stall event. The current study is aimed at validation of the developed flow solver using an industrial compressor database. Further, the current study is aimed at understanding the interaction between the stages with regards to pressure oscillations leading to stall.