Building upon the successes of the UDF® program in the 1980’s, open rotor designs for high flight speed efficiency and low community noise have been developed at GE in collaboration with NASA and the FAA. Targeting a narrow body aircraft with 0.78 cruise Mach number, the cost-share program leveraged computational fluid dynamics (CFD), computational aero-acoustics (CAA), and rig scale testing to generate designs that achieved significant noise reductions well beyond what was attained in the 1980’s while substantially retaining cruise performance. This paper presents overall propeller net efficiency and acoustic assessments of GE’s modern open rotor designs based on measured rig data and the progression of the technology from the 1980’s through the present. Also discussed are the effects of aft rotor clipping, inter-rotor spacing, and disk loading. This paper shows how the two-phase design and scale model wind tunnel test program allowed for test results of the first design phase to feed back into the second design phase, resulting in 2–3% improvement in overall propeller net efficiency than the best efficiency design of the 1980’s while nominally achieving 15–17 EPNdB noise margin to Chapter 4 (when projected to full scale for a prescribed aircraft trajectory and installation). Accounting for trades and near term advancements, such a propulsion system is projected to meet the goal of 26% fuel burn reduction relative to CFM56-7B powered narrow body aircraft.
A key objective of compressor rig tests is the identification of compressor stall boundary. A complementary goal is the identification of the stalling stage based on test data. This serves two purposes: 1) Validate the pre-test prediction of the stage loading distribution, and 2) identify the weak stages, should improvements in operating range be desired in subsequent design iterations. Typically the pertinent test data is in the form of static pressure measurements. Many engineers believe that a stalling stage is accompanied by a transient upstream pressure rise coupled with a downstream pressure loss. However, inter-stage dynamics may cloud the identification of the stalling stage. To this end, an analysis of inter-stage dynamics, immediately preceding the stall event, could provide an alternate assessment of the stalling stage. This work reviews existing stall models for studying compressor dynamics. The main focus of this work is to develop ability to capture inter-stage dynamics. A 3-state equation lumped Moore-Greitzer (MG3) model is widely used to study the dynamic compressor response during surge and rotating stall transients. However the evolution of MG3 model may not provide a suitable framework for the investigation of inter-stage dynamics. On the other hand, an unsteady time marching 1-D fluid dynamic model (e.g. similar to the DynTECC formulation which includes body forces), while unable to capture the rotating stall dynamics, is sufficient for this purpose. A numerical simulation has been developed 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. Through a controlled weakening of selected stages, the time evolution of these waves is related back to the stalling stage. It is found that the weakened stage is not necessarily the stalling stage as identified via the pressure rise and downstream pressure drop pattern.
Simulation of ice shed into a multistage axial compressor involves a coupled two phase flow of a continuous phase comprising of air and water vapor and a discrete phase with ice crystals and water droplets. A first principles based discrete phase model is formulated to capture the heat and mass transfer processes of ice flow in air. A quasi one-dimensional model is used to represent the continuous phase. An exchange of information at every time step between the two models leads to a coupled response that alters characteristics like temperature and pressure distributions across the compressor. However, an understanding of the impact of various assumptions used for modeling of the icing physics is imperative in order to establish the fidelity of the developed icing model, before its use in gas turbine engine ice ingestion studies. This paper describes the assumptions and semi-numerical models used in the coupled discrete-continuous phase flow numerical models. The input characteristics of the discrete phase related to the size and distribution of ice crystals, the assumed percentage of ice particles escaping through compressor bleed ports, simplifications associated with ice and droplet breakup on impact with compressor blades, moisture content affecting the dry air properties, are some of the factors that are variables in the icing study. The impact of these factors on the compressor flow dynamics is estimated through a parametric analysis.
The tip region of transonic blades in turbomachinery involves complex flow physics including shock wave/boundary layer interactions (SBLI). SBLI can lead to flow separation, transition from laminar to turbulent flow, and unsteadiness, which can affect the overall performance of the blade. In this paper, we present wall-resolved large eddy simulations (LES) of a transonic rotating cascade that is modeled after the tip region of a transonic diffusing blade. The calculations were performed using GENESIS, a high-order unstructured large eddy simulation solver. The convergence of the LES solution is assessed by varying the polynomial order of the solution from low to high order. LES simulations for a total of five operating conditions are presented, which cover the range of operation from unique incidence low operating line to stall and the associated shock wave/boundary layer interaction physics. The overall aerodynamics of the transonic passage airfoil are described based on the LES solutions as well as providing a detailed analysis of the boundary layer behavior. The changes in shock structure, boundary layer interaction physics, and associated losses with operating condition are highlighted. A low-frequency SBLI unsteadiness is observed in the cases where the boundary layer into the shock is laminar, and a scaling of the frequency is proposed. The scaling is based on the time scale of turbulent structures convecting from the shock to the trailing edge and acoustic disturbances then traveling back upstream.
Shock/boundary layer interactions (SBLI) are a fundamental fluid mechanics problem relevant in a wide range of applications including transonic rotors in turbomachinery. This paper uses wall-resolved large eddy simulation (LES) to examine the interaction of normal shocks with laminar and turbulent inflow boundary layers in transonic flow. The calculations were performed using GENESIS, a high-order, unstructured LES solver. The geometry created for this study is a transonic passage with a convergent-divergent nozzle that expands the flow to the desired Mach number upstream of the shock and then introduces constant radius curvature to simulate local airfoil camber. The Mach numbers in the divergent section of the transonic passage simulate single stage commercial fan blades. The results predicted with the LES calculations show significant differences between laminar and turbulent SBLI in terms of shock structure, boundary layer separation and transition, and aerodynamic losses. For laminar flow into the shock, significant flow separation and low-frequency unsteadiness occur, while for turbulent flow into the shock, both the boundary layer loss and the low-frequency unsteadiness are reduced.
The flow over modern turbofan blades is highly complex, with 3D, transonic and transitional boundary layer, shock-boundary layer interaction, secondary flow effects and complex wake evolution process. Traditional turbulent flow modeling techniques based on the linear Boussinesq approximation of the Reynolds stress and empirical boundary layer transition models have limited applicability and often fail to correctly predict the limits of flow separation and transitions. Large Eddy Simulation (LES) methods have considerably less modeling empiricism and hence are expected to predict such complex phenomenon better than most Reynolds Averaged Navier-Stokes (RANS) models. A methodology for performing Large Eddy Simulations is developed and applied to canonical blade cascade problems and a transonic 2D bump for experimental validation and extended to a modern 3D transonic turbofan. The LES solutions are compared to standard two-equation RANS predictions.
A new simple engineering parameter to evaluate the stability of multistage axial compressors has been derived. It is based on the stability analysis for a small circumferential disturbance imposed on the steady-state flow field. The analytical model assumes that the flow field is two dimensional and incompressible in the ducts between blade rows although the steady-state density is permitted to change across the blade rows. The resulting stall parameter contains terms that relate to the slope of the pressure rise characteristic of the blade rows and the inertia effects of the fluid in the blade rows and ducts. The parameter leads to the classical stability criteria based on the slope of the overall total to static pressure rise coefficient in the limit where constant density and constant blade rotational speed are assumed across the compressor. The proposed stall parameter has been calculated for three different multistage axial flow compressors, and the results indicate that the parameter has a strong correlation with the measured stability of the compressors. The good correlation with the test data demonstrates that the newly derived stall parameter captures much of the fundamental physics of instability inception in multistage compressors, and that it can be a good guideline for designers and engineers needing to evaluate the stability boundary of multistage machines.
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