Gas Turbine Performance Improvement by Retrofit of Advanced Technology

Development and operation of larger size gas turbines have demonstrated that higher turbine inlet temperature can be sustained due to advancement in material and cooling technology. After a feasibility study it was determined that modern available technology can be applied to existing previous generation of machines. These programs are identified as “The Performance Upgrade of Gas Turbine”. Amongst the significant benefits that can be realized by retrofitting state of art parts in existing machines are higher power and more durable parts. This paper discusses various programs that are currently offered and implementation technique of upgrading the machines. A recent example is also presented. These unique programs are particularly attractive at the time of overall life consumption of the initial set of hot parts. At that point in an operating gas turbine it will be beneficial to retrofit the latest configuration parts to realize the performance improvements.

Field Test of the W501D, a 100 MW Combustion Turbine

The field test of the current production W501D combustion turbine shows the engine to be an outstanding success. The engine meets power and exceeds efficiency expectations. Metal temperatures are at or below expected levels throughout the engine. In particular the test data demonstrates low temperatures provided by improved design concepts for the combustion chamber and first stage turbine vane, which are critical components because of their severe environment. Other design improvements for performance and reliability were also verified during the test.

Heat Transfer Characteristics for Jet Array Impingement With Initial Crossflow

Two-dimensional arrays of circular air jets impinging on a heat transfer surface parallel to the jet orifice plate are considered. The jet flow, after impingement, is constrained to exit in a single direction along the channel formed by the jet orifice plate and the heat transfer surface. In addition to the crossflow which originates from the jets following impingement, an initial crossflow is present which approaches the array through an upstream extension of the channel. The temperature of the initial crossflow air may differ from the jet air temperature. The configurations considered are intended to model the impingement cooled midchord region of gas turbine airfoils in cases where an initial crossflow is also present. Nusselt numbers and dimensionless adiabatic wall temperatures resolved to one streamwise jet hole spacing were experimentally determined for ratios of the initial crossflow rate to the total jet flow rate ranging from zero to unity. These are presented and discussed relative to the flow and geometric parameters.

Film Cooling With Steam Injection Through Three Staggered Rows of Inclined Holes Over a Straight Airfoil

Experiments have been performed to investigate the film cooling characteristics of steam injection through three staggered rows of 60 degree inclined holes over a straight aluminum airfoil with a circular leading edge. The axial distance between cooling hole rows is five hole diameters. The lateral distance is four hole diameters. Data have been taken for the local film cooling effectiveness of steam and air with blowing rates varying from 0.3 to 1.8. It shows that at small blowing rates, steam has an effectiveness up to 2.5 times greater than air; but, as the blowing rates are increased, the difference between the steam and air effectiveness is gradually decreased. It also shows that the steam effectiveness is less dependent upon the blowing rates than is the air effectiveness. The results generally support the previous analytical prediction.

Measured and Calculated Wall Temperatures on Air-Cooled Turbine Vanes With Boundary Layer Transition

Convection cooled turbine vane metal wall temperatures experimentally obtained in a hot cascade for a given one-vane design were compared with wall temperatures calculated with TACT1 and STAN5 computer codes which incorporated various models for predicting laminar-to-turbulent boundary layer transition. Favorable comparisons on both vane surfaces were obtained at high Reynolds number with only one of these transition models. When other models were used, temperature differences between calculated and experimental data obtained at the high Reynolds number were as much as 14 percent in the separation bubble region of the pressure surface. On the suction surface and at lower Reynolds number, predictions and data unsatisfactorily differed by as much as 22 percent. Temperature differences of this magnitude can represent orders of magnitude error in blade life prediction.

Convective Heat Transfer on Flat Plate at Very High Temperature and Pressure Gradient

This paper describes heat transfer measurements on flat plates for various free-stream temperatures, pressure gradients, Reynolds numbers and Mach numbers anticipated in advanced gas turbines. A shock tube generated the high temperature and pressure air flow, and a variable geometry test section was used to produce inlet flow Mach number of 0.11 and accelerate the flow over the plate to sonic velocity. Thin platinum film heat gages recorded the local surface heat flux for laminar, transition, and turbulent boundary layers. The free-stream temperatures varied from 630°R (350°K) to 4700°R (2611°K) for a Tw/Tr,g temperature ratio of 0.82 to 0.12 and Mach numbers from 0.11 to 1.90. The Reynolds number based on the model length was varied from 8.8 × 102 to 5 × 106. The experimental heat flux data were correlated with the laminar and turbulent theories and the transition phenomenon was examined.

Comparison of Predicted and Experimental External Heat Transfer Around a Film Cooled Cylinder in Crossflow

Calculations were made of the film cooling provided by rows of holes around the circumference of a cylinder in crossflow and the results were compared to experimental data obtained from a NASA grant to Purdue University. The calculations and experimental data were for conditions that simulate most of those that are typical of air cooled turbine vane leading edges. Injection was from single and multiple rows of holes located at different angular locations from the stagnation line. The holes in the rows were angled normal to the flow direction and at a 25 degree angle to the cylinder wall. The calculations and experimental data were for several constant values of blowing ratios for all rows and for different blowing ratios for each row, representing a simulation of a common coolant plenum supply to multiple rows of holes. The calculations were made using a finite difference boundary layer code, STAN5, developed under NASA contract with Stanford University and modified at the NASA Lewis Research Center. Contrary to initial expectations that injection would trip the boundary layer flow into the turbulent regime, the results indicated that the high free stream acceleration apparently kept the flow laminar for holes in the first 45 degrees past stagnation. The trend in Stanton number reduction due to coolant injection was predicted with generally good agreement at the lower blowing rates, but for multiple rows of holes, agreement was poor beyond the first row.

Heat Transfer Experiments in High Aspect Ratio Rectangular Channel With Epoxied Short Pin Fins

In an attempt to reduce the cost of testing many configurations of short pin fins in a rectangular channel, a technique has been identified whereby the pins are epoxied to the end wall and can be easily removed to form a new configuration at the end of a test. Analytical and experimental results indicate that the temperature drop across a thin layer of epoxy (∼.005–.006 cm) (K = 22.5 W/m°C) with copper pin and endwalls was less than 1% of the heat transfer surface temperature. The technique was then used to test 4 pin fin configurations of height to diameter ratio of about unity. The heat transfer results showed excellent agreement with earlier published data, thus confirming the validity of this technique.

A Small Engine High Temperature Core Research Programme

A general description of a small engine high temperature research vehicle is presented along with other aspects of the associated programme directed primarily towards the investigation of an air cooled turbine system. The general philosophy and style of the programme is discussed and comparisons made with alternative approaches. The instrumentation used is detailed along with a brief review of the test facility and data retrieval system. The presentation of engine design features and results is limited due to the classified nature of the work. The project is reviewed in the light of the 5 years test experience gained.

New Radial Engine Technology From Kongsburg

The first engine of a new family of high performance, rugged radial turbines is presented. The new single shaft 1500 kW site rated KG3 is built upon extensive experience from the field proven KG2 of the same nominal power. The KG3 is being developed both as a simple cycle and recuperated engine providing specific fuel consumptions in the range of 0.29 to 0.19 kg/kW·hr (0.47 to 0.32 lbs/hp·hr). The new Kongsberg engine makes use of a high specific speed high pressure ratio centrifugal compressor combined with a very high tip speed uncooled radial turbine to obtain optimized aerodynamic matching. Several novel design features are described.