GAS TURBINE ENGINE COMPRESSOR-COMBUSTOR SIMULATION

The superior output power-to-weight ratio of the gas turbine engine has made it the mainstay of modern aviation propulsion for both civilian and military applications. The capability of producing high output power at low engine weight has also made it a common propulsion system in both marine and armored military ground applications. In addition, gas turbine engines are used for power generation as either stand-alone systems or coupled with steam turbines to power electrical generators and also to power rotating machinery such as oil pumps in large pipeline systems. To gain an edge in the marketplace, companies are continually striving to improve the efficiency and performance of their engines. Ideally, this would involve a complete engine simulation, providing the capability for the flow fields through the various components to interact. Thus, the actual physical processes would be simulated, including the effects of three-dimensional, unsteady, turbulent viscous reacting flows and their interaction with the engine structural components. Unfortunately, a complete engine simulation model requires vast computational resources. As a result, simplified models have been developed that primarily focus on the simulation of steady flow phenomena in the individual engine components. Thus, individual steady flow performance models have been developed for each component and their performance is predicted independently.

Gas Turbine Engine Steady Performance
The surge and choke lines bound the operating range of a gas turbine engine on the compressor aerodynamic performance map, Figure 1.

Figure 1. Typical performance map.

To assure compressor stability during engine operation, an engine is designed with a surge margin. This entails assuring that the operating point remains a specified distance from the surge line on the performance map. Large surge margins are employed due to transient conditions that move the compressor operating point closer to the surge line. However, a large surge margin that places the compressor operating line far from the surge line can preclude operation at the peak pressure rise or maximum efficiency region. Also, the increase in operational range results in additional flexibility for matching the compressor with the other gas turbine engine components.

The term surge line is actually a misnomer as two types of instability can develop: surge and rotating stall. Surge is a global axisymmetric oscillation of the flow through the compressor that can include reverse flow during a portion of the surge cycle. These oscillations can result in severe damage to the mechanical components of the engine from the unsteady thrust load or the ingestion of combustion gases into the compressor and engine inlet. In a severe surge cycle, the reversed flow through the compressor can extinguish combustion, resulting in a "flame out" or total loss of engine power.

Rotating stall is a local flow deficit that rotates around the compressor annulus. This flow deficit, or cell, is a region in which the local mass flow is near zero. The rotating stall may consist of one or multiple cells that rotate around the compressor at an angular speed which is a fraction of the rotor speed. This instability results in a loss of compressor performance that may require the shut down of the engine.

Transient Performance
The performance of a gas turbine engine can differ significantly from that predicted from such independent steady flow models because of unsteady interactions that occur. The consequences of these dynamics can be quite dramatic, including the unexpected crossing of the compressor surge line while transitioning between engine operating points, as depicted in Figure 2. The unexpected crossing of the compressor surge line during engine transients results in a complex dynamic interaction between the engine components driven by rotating stall and surge.

Figure 2. Transient performance can cause the engine to cross the surge line.

This unsteady engine operation produces extreme loading for the turbomachinery blading, resulting in high cycle fatigue (HCF) failures, with surge and rotating stall resulting in dangerous flow induced blade vibrations due to the rapid loading and unloading of the blading.

To address the various issues associated with the transient performance of a gas turbine engine, an advanced simulation will be implemented that models the inlet, fan, compressor and combustor. Thus, this simulation will capture the important compressor-combustor interactions that occur during engine transients, including rotating stall and surge, and their affect on blade row durability. This simulation will provide the capability for the flow fields through the various components to interact. Thus, the actual physical processes will be simulated, including the effects of three-dimensional, unsteady, turbulent viscous reacting flows and their interaction with the engine structural components.

Turbomachinery Aerodynamics & Blade Row Fluid-Structure Interactions
Classical Approach: The primary mechanism of blade failure is fatigue caused by vibrations at levels exceeding material endurance limits. The classical approach is to predict the amplitude responses and stresses at the resonant speeds with the fluid and structure are modeled separately, i.e., they are not coupled. They are then coupled by specifying the kinematic boundary conditions at the fluid-structure boundary. Thus, the unsteady aerodynamic forces acting on the blading are predicted with the motion of the structure as a boundary condition. Such unsteady flow models assume that (1) the unsteadiness is a small perturbation from the steady flow and (2) the airfoil motion is specified.

Unfortunately, this uncoupling of a truly coupled problem introduces error as the blade row unsteady aerodynamic loading is dependent on the specified motion of the blade. Thus, instead of utilizing separate fluid and structural models, a coupled interacting fluid-structures analysis is needed. In this regard, the finite element code ALE3D (Arbitrary Lagrange/Eulerian 3D Code System) developed at Lawrence Livermore National Lab is a finite element code whose formulation is general enough to model both fluids and structures in a Lagrangian, an Eulerian, or an intermediate perspective.

Research Plan: A next generation coupled fluid-structure interaction model for turbomachine blade rows will be developed. This will be accomplished by extending ALE3D to address the unsteady aerodynamics of turbomachine blade rows. Inputs include the blade row geometry, the inlet flow field and the mass flow, specified by means of the blade row exit pressure. The output includes the blade row inlet flow field, iterated with the input to meet the mass flow requirement, and the blade row flow field, including the exit flow field. Note that this exit flow field is the combustor inlet flow field.

The issues associated with transient gas turbine engine performance will then be addressed. This will initially accomplished by applying a time-varying exit flow boundary condition to simulate rotating stall and surge conditions, with the resulting blade row unsteady loading predicted.

Combustor Flow Field
The gas turbine engine combustor increases the enthalpy of the working fluid by oxidization of fuel and the subsequent dilution of the resulting products with additional air until temperatures acceptable to the turbine are achieved. CFD gas turbine combustor modeling has generally been limited to isolated parts of the combustion system. Most models include only the reacting flow inside the combustor liner with assumed profiles and flow spits at the various liner inlets. A CFD calculation for the unsteady flow through a complete annulus combustor - from the compressor diffuser exit to the turbine inlet - is needed. The model should include an airblast fuel nozzle, dome and liner walls with dilution holes and cooling louvers.

Research Plan: The combustor will be modeled by the KIVA code modified for the gas turbine engine combustors. Inputs include the combustor geometry, the inlet flow field an the mass flow, specified by means of the combustor exit pressure. The output includes the combustor flow field and performance, iterated with the input to meet the mass flow requirement, and the combustor flow field, including the inlet flow field. Note that this inlet flow field is the blade row exit flow field. The steady performance of a combustor will be predicted first. The issues associated with transient gas turbine engine performance will then be addressed. This will initially be accomplished by applying time-varying inlet flow conditions to the combustor, with the resulting time-varying combustor performance predicted.

Simulation of Gas Turbine Engine Transient Performance
The advanced simulation of transient gas turbine engine performance will capture the important compressor-combustor interactions that occur during engine transients, including rotating stall and surge, and their affect on blade row durability. This will be accomplished by coupling the unsteady flow field analysis ALE3D and the combustor model KIVA. This coupled simulation is depicted in Figure 3. The mass flow requirement is met through the combustor exit pressure condition. The blade row and combustor simulations are coupled through the ALE3D predicted blade row exit flow field that is the combustor KIVA input and the KIVA predicted combustor inlet flow field which is the blade row exit pressure.