A Three-Dimensional Viscous Transonic Inverse Design Method

The development and application of a three-dimensional inverse methodology is presented for the design of turbomachinery blades. The method is based on the mass-averaged swirl, rV~θ distribution and computes the necessary blade changes directly from the discrepancies between the target and initial distributions. The flow solution and blade modification converge simultaneously giving the final blade geometry and the corresponding steady state flow solution. The flow analysis is performed using a cell-vertex finite volume time-marching algorithm employing the multistage Runge-Kutta integrator in conjunction with accelerating techniques (local time stepping and grid sequencing). To account for viscous effects, dissipative forces are included in the Euler solver using the log-law and mixing length models. The design method can be used with any existing solver solving the same flow equations without any modifications to the blade surface wall boundary condition. Validation of the method has been carried out using a transonic annular turbine nozzle and NASA rotor 67. Finally, the method is demonstrated on the re-design of the blades.

LM2500 Reliability Improvements for United States Navy Applications

The United States Navy uses the General Electric LM2500 gas turbine engine for main propulsion on its newest surface combatants including the OLIVER HAZARD PERRY (FFG 7) class frigates, SPRUANCE (DD 963) class destroyers, TICONDEROGA (CG 47) class cruisers, ARLIEGH BURKE (DDG 51) class destroyers and SUPPLY (AOE 6) class oilers. Currently, the Navy operates a fleet of over 400 LM2500 gas turbine engines. This paper discusses the ongoing efforts to characterize the availability of the engines aboard ship and pinpoint systems/components that have significant impact on engine reliability. In addition, the program plan to upgrade the LM2500’s standard configuration to improve reliability is delineated.

Aircraft Propulsion System Accident Causal Analysis

A causal analysis of aviation accidents by engine type is presented. The analysis employs a top-down methodology that performs a detailed analysis of the causes and factors cited in accident reports to develop a “fingerprint” profile for each engine type. This is followed by an in-depth analysis of each fingerprint that produces a sequential breakdown. Analysis results of National Transportation Safety Board (NTSB) accidents, both fatal and non-fatal, that occurred during the time period of 1990–1998 are presented. Each data set is comprised of all accidents that involved aircraft with the following engine types: turbofan, turbojet, turboprop, and turboshaft (includes turbine helicopters). During this time frame there were 1461 accidents involving turbine powered aircraft; 306 of these involved propulsion malfunctions and/ or failures. Analyses are performed to investigate the sequential relationships between propulsion system malfunctions or failures with other causes and factors for each engine type. Other malfunctions or events prominent within each data set are also analyzed. Significant trends are identified. The results from this study can be used to identify areas for future research into intervention, prevention, and mitigation strategies.

Effects of Tip Clearance on Hot Streak Migration in a High-Subsonic Single-Stage Turbine

Experimental data have shown that combustor temperature non-uniformities can lead to the excessive heating of first-stage rotor blades in turbines. This heating of the rotor blades can lead to thermal fatigue and degrade turbine performance. The results of recent studies have shown that variations in the circumferential location, or clocking, of the first-stage vane airfoils can be used to minimize the adverse effects of the hot streaks due to the hot fluid mixing with the cooler fluid contained in the vane wake. In addition, the effects of the hot streak/airfoil count ratio on the heating patterns of turbine airfoils have been quantified. In the present investigation, three-dimensional unsteady Navier-Stokes simulations have been performed for a single-stage high-pressure turbine geometry operating in high subsonic flow to study the effects of tip clearance on hot streak migration. Baseline simulations were initially performed without hot streaks to compare with the experimental data. Two simulations were then performed with a superimposed combustor hot streak; in the first the tip clearance was set at the experimental value, while in the second the rotor was allowed to scrape along the outer case (i.e., the limit as the tip clearance goes to zero). The predicted results for the baseline simulations show good agreement with the available experimental data. The simulations with the hot streak indicate that the tip clearance increases the radial spreading of the hot fluid, and increases the integrated rotor surface temperature compared to the case without tip clearance.

A Test-Based Methodology for A Priori Selection of Gain/Phase Relationships in Proportional, Phase-Shifting Control of Combustion Instabilities

Feedback control system design, for general single-in-single-out (SISO) applications, requires accurate knowledge of the loop transfer function. Active combustion control design is usually implemented using such SISO architectures, but is quite challenging because the thermoacoustic response results from a relatively unknown, self-excited system and nonlinear processes that must be understood before learning the gain/phase relationship of the system precisely at the instability frequency. However, recent experiments have shown that it is possible to obtain accurate measurements of the relevant loop transfer (frequency response) functions at frequencies adjacent to the instability frequency. Using a simple tube combustor, operating with a premixed, gaseous, burner-stabilized flame, the loop frequency response measurements have been used to develop a methodology that leads to ‘test-based predictions’ of the absolute phase settings and ‘best’ gain settings for a proportional, phase-shifting controller commanding an acoustic actuator in the combustor. The contributions of this methodology are twofold. First, it means that a manual search for the required phase setting of the controller is no longer necessary. In fact, this technique allows the absolute value of controller phase to be determined without running the controller. To the authors’ knowledge, this has not been previously reported in the literature. In addition, the ‘best’ gain setting of the controller, based on this new design approach, can be defined as one that eliminates or reduces the limit cycle amplitude as much as possible within the constraint of avoiding generation of any controller-induced instabilities. (This refers to the generation of ‘new’ peaks in the controlled acoustic pressure spectrum.) It is shown that this tradeoff in limit cycle suppression and avoidance of controller-induced instabilities is a manifestation of the well-known tradeoff in the sensitivity/complementary sensitivity function for feedback control solutions. The focus of this article is limited to the presentation of the design method and does not discuss the detailed nonlinear phenomena that must be understood to determine the optimal gain/phase settings at the limit cycle frequency for a real (versus theoretical) combustor system. A companion paper describes how the proposed design method can be used to generate an AI controller that maintains stabilizing control for a range of changing operating conditions.

Tip Clearance Flow Induced Endwall Boundary Layer Separation in a Single–Stage Axial–Flow Low–Speed Compressor

The paper describes an investigation of the overtip end-wall flow in a single–stage axial–flow low–speed compressor utilizing an oil flow technique and a periodic multisampling pressure measurement technique. Representative oil flow pictures and ensemble averaged casingwall pressure distributions with standard deviations — supplemented by selected endwall oil flow pictures from a corresponding 2D compressor cascade — are shown and carefully analysed. The results enable the key features of the overtip endwall flow to be identified and changes with flow rate — or inlet angle — to be determined.

Gas Turbine Control System Integration for Royal Caribbean Millennium Class Cruise Liner

Traditionally commercial marine propulsion needs have been met with direct drive reciprocating prime movers. In order to increase efficiency, simplify installation and maintenance accessibility, and increase cargo / passenger capacity; indirect electric drive gas and steam turbine combined cycle prime movers are being introduced to marine propulsion systems. One such application is the Royal Caribbean Cruise Line (RCCL) Millennium Class ship. This commercial vessel has two aero-derivative gas turbine generator sets with a single waste heat recovery steam turbine generator set. Each is controlled by independent microprocessor based digital control systems. This paper addresses only the gas turbine control system architecture and the unique safety and dynamic features that are integrated into the control system for this application.

One-Sided Control of Surge in a Centrifugal Compressor System

Active surge control is studied in a centrifugal compressor system. To stabilize surge, a bleed valve is applied which is nominally closed and only opens to stabilize the system around the desired operating point. This bleed valve is controlled using a linear output feedback controller based on plenum pressure measurements. In this study, the practical limitations of this one-sided surge controller are examined. Experiments show that the performance of the controller is affected by measurement noise and the desired equilibrium point in the controller, which is not accurately known in the unstable operating region. Fully-developed surge is stabilized with relatively small stationary bleed valve mass flow using the proposed controller in combination with a small band-pass filter. Measurements are compared with the results of the Greitzer model. This model shows qualitative agreement with experiments. For the examined compressor system, a 3.5% extension of the stable operating region is obtained using the one-sided surge controller.

On the Modeling of Tip Leakage Flow in Axial Turbine Blade Rows

The starting point of this paper is an established turbine tip leakage loss model based on energy considerations. The model requires a discharge coefficient as an empirical input. The discharge coefficient is the ratio of the actual to the theoretical tip gap mass flow rate, The nondimensional parameters influencing the discharge coefficient are determined by a dimensional analysis. These parameters are: gap width to length ratio, end wall speed to gap flow velocity ratio and gap Reynolds number. Ranges for these parameters, valid for typical turbine tip gap situations, are presented. The numerical investigation of the turbulent flow in a plane perpendicular to the blade chord line supplies the discharge coefficient versus the nondimensional gap width. Depending on the gap width to length ratio, various degrees of mixing of the flow downstream of the vena contracta can be detected. Based on these observations, a simple tip gap flow model is presented. The discharge coefficients computed by this model are compared with the numerical results as well as with experimental values from the literature. Finally, the model is used to calculate the discharge coefficients of improved tip gap geometries (squealers, winglets).

The Effect of Airfoil Scaling on the Predicted Unsteady Loading on the Blade of a 1 and 1/2 Stage Transonic Turbine and a Comparison With Experimental Results

In this study, two time-accurate Navier-Stokes analyses were obtained to predict the first-vane/first-blade interaction in a 1 and 1/2-stage turbine rig for comparison with measurements. In the first computation, airfoil scaling was applied to the turbine blade to achieve periodicity in the circumferential direction while modeling 1/18 of the annulus. In the second, 1/4 of the wheel was modeled without the use of airfoil scaling. For both simulations the predicted unsteady pressures on the blade were similar in terms of time-averaged pressure distributions and peak-peak unsteady pressure envelopes. However, closer inspection of the predictions in the frequency domain revealed significant differences in the magnitudes of unsteadiness at twice vane-passing frequency (and the vane-passing frequency itself, to a lesser extent). The results of both computations were compared to measurements of the vane-blade interaction in a full-scale turbine rig representative of an early design iteration of the PW6000 engine. These measurements were made in the short-duration turbine-test facility at The Ohio State University Gas Turbine Laboratory. The experimentally determined, time-resolved pressures were in good agreement with those predicted with the 1/4-wheel simulation.