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.

The Potential Impact of Utilizing Advanced Engine Technology for a Combat Capable Unmanned Air Vehicle (UAV)

This paper evaluates the potential impact of utilizing advanced engine technology for a limited life, combat capable, unmanned air vehicle (UAV) application. A study was conducted to define payoffs in terms of mission capability and system level life cycle costs (LCC) associated with implementing three different engine development approaches into a combat capable UAV design. The three different approaches considered were: a new, advanced technology engine; an existing (off-the-shelf) engine; and a derivative of an existing engine with limited technology insertion. A detailed vehicle configuration design was developed to conduct this assessment, including a low observable (LO), highly integrated engine/airframe layout for survivability and mission adaptable considerations. The vehicle is designed with multi-role mission capability such as suppression of enemy air defense (SEAD), close air support (CAS), and battlefield air interdiction (BAI). A system level performance comparison is assessed with the three different engine approaches, specifically for the SEAD-type mission. For the cost analysis, the multi-role mission capability is reflected in the overall assumptions such as in the number of aircraft needed to meet the mission requirements. A system level assessment such as in this study is essential in determining whether the additional costs associated with the development of a new, advanced engine is worth the investment. The results of this study suggest that advanced engine technology insertion can provide significant benefits in terms of mission range capability, vehicle weight/size, and overall life cycle costs versus an existing engine.

Analysis of the Mixing Plane Interface Between Stator and Rotor of a Transonic Axial Turbine Stage

A three-dimensional transonic turbine stage is computed by means of a numerical simulation tool. The simulation accounts for the coolant ejection from the stator blade and for the tip leakage of the rotor blade. The stator and rotor rows interact via a mixing plane, which allows the stage to be computed in a steady manner. The analysis is focused on the matching of the stator and rotor mass flow rates. The computations proved that the mixing plane approach allows the stator and rotor mass flow rates to be balanced with a careful choice of the stator-rotor static pressure interface. At the same time, the pitch averaged distribution of the transported quantities at the interface for the stator and rotor may differ slightly, together with the value of the static pressure at the hub.

Investigation of the Unsteady Rotor Flow Field in a Single HP Turbine Stage

Within a European project a high-pressure turbine stage was investigated at DLR, Göttingen. The investigations consisted primarily of experiments carried out in the windtunnel for Rotating Cascades (RGG), but some numerical work was also performed. Detailed measurements were carried out at mid section of a turbine rotor using a Laser-2-Focus device which served as a velocimeter measuring 2D-velocity vectors and turbulence quantities and as a tool to determine the concentration of coolant ejected at the trailing edge of the stator blades. The measurement of coolant concentration downstream of the stator and inside the rotor provided a detailed picture of the stator wake development and its interaction with the moving rotor. Axial measurement locations reached from the stator exit through the rotor to a downstream measurement plane. Measurement results are presented as instantaneous flow values. Unsteady flow vectors and turbulence intensities could be related at 16 time instants representing one rotor blade passsing period to the wake development made visible by the coolant concentration. The measured unsteady flow vectors and unsteady pressures, measured with semi-conductor pressure transducers, are compared with results from a numerical calculation using the Navier-Stokes code “TRACE-U” which allows the computation of the unsteady flow field. The measured steady and unsteady flow quantities served to validate the Navier-Stokes code. A comparison of the wake entropy trajectories outside the blade boundary layers and at the wall gives an impression of the lag between the arrival time of the wake in the freestream near the blade surface and the time the boundary layer quantities at the blade surface itself are affected.

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.