Gas turbine combustion chambers were traditionally designed through trial and error which was unfortunately a time-consuming and expensive process. The development of computers, however, contributed a great deal to the development of combustion chambers, enabling one to model such systems more accurately in less time. Traditionally, preliminary combustor designs were conducted with the use of one-dimensional codes to assist in the prediction of flow distributions and pressure losses across the combustion chamber mainly due to their rapid execution times and ease of use. The results are generally used as boundary conditions in three- dimensional models to predict the internal flow field of the combustor. More recent studies solve the entire flow field from prediffuser to combustor exit. This approach is, however, a computationally expensive procedure and can only be used if adequate computer resources are available. The purpose of this study is two-fold; (1) to develop a one-dimensional incompressible code, incorporating an empirical-based combustion model, to assist a one-dimensional network solver in predicting flow- and temperature distributions, as well as pressure losses. This is done due to the lack of a combustion model in the network solver that was used. An incompressible solution of flow splits, pressure losses, and temperature distributions is also obtained and compared with the compressible solution obtained by the network solver; (2) to utilise the data, obtained from the network solver, as boundary conditions to a three-dimensional numerical model to investigate possible modifications to the dome wall of a standard T56 combustion chamber. A numerical base case model is validated against experimental exit temperature data, and based upon that comparison, the remaining numerical models are compared with the numerical base case. The effect of the modification on the dome wall temperature is therefore apparent when the modified numerical model is compared with the numerical base case. A second empirical code was developed to design the geometry of axial straight vane swirlers with different swirl angles. To maintain overall engine efficiency, the pressure loss that was determined from the network analysis, of the base case model, is used during the design of the different swirlers. The pressure loss across the modified combustion chamber will therefore remain similar to that of the original design. Hence, to maintain a constant pressure loss across the modified combustion chambers, the network solver is used to determine how many existing hole features should be closed for the pressure loss to remain similar. The hole features are closed, virtually, in such a manner as not to influence the equivalence ratio in each zone significantly, therefore maintaining combustion performance similar to that of the original design. Although the equivalence ratios in each combustion zone will be more or less unaffected, the addition of a swirler will influence the emission levels obtained from the system due to enhanced air-fuel mixing. A purely numerical parametric analysis was conducted to investigate the influence of different swirler geometries on the dome wall temperature while maintaining an acceptable exit temperature distribution. The data is compared against the data obtained from an experimentally validated base case model. The investigation concerns the replacement of the existing splash-cooling devices on the dome wall with that of a single swirler. A number of swirler parameters such as blade angle, mass flow rate, and number of blades were varied during the study, investigating its influence on the dome wall temperature distribution. Results showed that the swirlers with approximately the same mass flow as the existing splash-cooling devices had almost no impact on the dome wall temperatures but maintained the exit temperature profile. An investigation of swirlers with an increased mass flow rate was also done and results showed that these swirlers had a better impact on the dome wall temperatures. However, due to the increased mass flow rate, stable combustion is not guaranteed since the air/fuel ratio in the primary combustion zone was altered. The conclusion that was drawn from the study, was that by simply adding an axial air swirler might reduce high-temperature gradients on the dome but will not guarantee stable combustion during off-design operating conditions. Therefore, a complete new hole layout design might be necessary to ensure good combustion performance across a wide operating range. / Dissertation (MEng (Mechanical))--University of Pretoria, 2008. / Mechanical and Aeronautical Engineering / MEng / unrestricted

No description available.

This thesis starts out with a brief description of gas turbine engines and information on railroad locomotives being the gas-turbine electric locomotives with some comparison of the diesel-electric locomotives in the introduction. Section 1.1 is the research problem looking at the older gas turbine electric locomotives in the 1950’s that ran on the rail and the problems they suffered. In section 1.2 titled the purpose of the study takes a look at newer gas turbine locomotives that were being consider or has been built with improvements since the 1950’s. The objective of the study being section 1.3 looks at the advantages of new gas turbines engines. Section 1.4 titled the research questions discusses better materials and methods of gas turbine engines. Chapter 2 is the literature review looking at the fuel oil specifications being number 4, number 5, and number 6. This chapter also talks about the used of distillates, types of distillates, composition of distillates, specifications for distillates, residual fuel oil and fuel oil quality dealing with the firing of gas turbine engines. Section 2.3 of chapter 2 being titled power generation looks at power plant gas-turbine engines and the power they produce. Chapter 3, titled the proposed methodology looks at setting up an experiment using a gas-turbine engine and a diesel-electric engine to compare the advantages of along with the disadvantages. Section 3.1 is titled data collected, within this section is discussion on the data collected from the experiment and improvements that could be made to the gas turbine engines. The end of chapter 3, section 3.2 titled data analyzing, talks about possible the results collected, calculations done, improvements made and rerunning another experiment with the improvements made. Chapter 4 discuss the types of materials using in building the compressor and turbine blades. Last, but not least is chapter 5 which discusses the actual experiment using the gas turbine simulator for aircrafts and how to apply it to the railroad locomotives. After the conclusion which discusses the results, is the appendix a being gas tables, appendix b being trial run 1 and appendix c being trial run 2.

Four series of tests were performed in an accelerated deposition test facility to study the independent effects of particle size, gas temperature, and metal temperature on ash deposits from two candidate power turbine synfuels. The facility matches the gas temperature and velocity of modern first stage high pressure turbine vanes while accelerating the deposition process. This is done by matching the net throughput of particulate out of the combustor with that experienced by a modern power turbine. In the first series of tests, four different size particles were studied by seeding a natural-gas combustor with finely-ground coal ash particulate. The entrained ash particles were accelerated to a combustor exit flow Mach number of 0.25 before impinging on a thermal barrier coated (TBC) target button at 1183°C. Particle size was found to have a significant effect on capture efficiency with larger particles causing considerable TBC spallation during a 4-hour accelerated test. In the second series of tests, different gas temperatures were studied while the facility maintained a constant exit velocity of 170 m/s (Mach=0.23-0.26). Coal ash with a mass mean diameter of 3 μm was used. Particle deposition rate was found to decrease with decreasing gas temperature. The threshold gas temperature for deposition was approximately 960°C. In the third and fourth test series impingement cooling was applied to the backside of the target button to simulate internal vane cooling. Ground coal and petcoke ash particulates were used for the two tests, respectively. Capture efficiency was reduced with increasing mass flow of coolant air. However, at low levels of cooling the deposits attached more tenaciously to the TBC layer. Post exposure analyses of the third and fourth test series (scanning electron microscopy and x-ray spectroscopy) show decreasing deposit thickness with increased cooling levels. Implications for the power generation goal of fuel flexibility are discussed.

Wind turbine aerodynamics are characterized by several coefficients, most notably the thrust, power, and axial induction, which is the fractional decrease of the free stream wind speed to the rotor plane. Current turbine designs aim to maximize these coefficients to reach what is generally considered to be maximum aerodynamic efficiency. Such rotors are referred to as a Betz-optimal rotor. This thesis examines a new method called “Low-Induction Rotors (LIR)” for increasing aerodynamic efficiency by decreasing the thrust loading of the blade. A family of low induction rotors (LIRs) can be derived from a reference wind turbine (RWT) by using the root bending moment as a constraint. Using the root bending moment of the RWT and imposed loading, new rotor lengths are derived. The family of low induction rotors are characterized by lower thrust loading across the blade. Prandtl’s bell shaped loading distribution was used to define the distribution of spanwise thrust since it is better fit for long-thin airfoils. Momentum Theory and Blade Element and Momentum theory were used to solve for the rotor power coefficient in two different ways, either including or excluding tip-losses. The family of rotors was then analyzed to determine power output. It was found that more power was produced as the rotor length increases and thrust loading decreased. The National Renewable Energy Lab’s AeroDyn software was used to conduct cp-λ sweeps on 6 selected rotors (128m, 137m, 147m, 156m, 167m, and 177m) to determine how power output was affected by changes in wind speed. The cp-λ analysis showed that the longer rotors with lower induction were less sensitive to changes in wind speed. The low induction rotors minimized a change in the coefficient of power as the pitch and tip speed ratio were changed. Low induction rotors are a promising field of wind energy, while maintaining the forces are the turbine hub, longer rotors with lower aerodynamic loading can be used to generate a more stable power supply.

No description available.

The objective of this investigation was to improve the thermal environment for a turbine vane through reduction of passage secondary flows. This was accomplished by modifying the vane/endwall junction to include a leading edge fillet. The problem approach was to integrate optimization methods with computational fluid dynamics to optimize the fillet design. The resulting leading edge fillet was then tested in a large-scale, low speed cascade to verify thermal performance. A combustor simulator located upstream of the cascade was used to generate realistic inlet conditions for the turbine vane. Both computational and experimental results underscore the importance of properly modeling the inlet conditions to the turbine. Results of the computational optimization process indicate that significant reductions in adiabatic wall temperature can be achieved with a leading edge fillet. While the intent of the initial fillet design was to improve the thermal environment for the vane endwall, computational results also indicate thermal benefit to the vane surfaces. Flow and thermal field results show that a fillet can enhance coolant effectiveness, prevent formation of the leading edge horseshoe vortex, and preclude full development of a passage vortex. In experimental testing, four cascade inlet conditions were investigated to evaluate the effectiveness of the fillet in reducing endwall temperature levels. Two tested conditions featured a flush combustor/cascade interface, while the remaining two included coolant injection through a backward-facing slot. With the flush interface, fillet thermal performance was evaluated for two inlet total pressure profiles. For the design profile, the fillet had a positive impact on the endwall temperature distribution as well as on the passage thermal field. For the off-design profile, the fillet was observed to have a slightly detrimental impact on the endwall adiabatic temperature distribution; however, passage thermal field results indicate a thermal benefit for the vane suction surface. With the backward-facing slot, thermal tests were conducted for two slot coolant flow rates. For both slot flow rates, the fillet improved endwall thermal protection and prevented coolant lift-off. While increasing the flow rate of slot coolant enhanced endwall effectiveness, fillet thermal performance was similar for the two slot flow rates. / Ph. D.

Cooling combustor chambers for gas turbine engines is challenging because of the complex flow fields inherent to this engine component. This complexity, in part, arises from the interaction of high momentum dilution jets required to mix the fuel with effusion film cooling jets that are intended to cool the combustor walls. The dilution and film cooling flow have different performance criteria, often resulting in conflicting flow mechanisms. The purpose of this study is to evaluate the influence that the dilution jets have on the film cooling effectiveness and how the flow and thermal patterns in the cooling layer are affected by both the dilution flow and the closely spaced film cooling holes. This study also intends to characterize the development of the flow field created by effusion cooling injection without dilution injection. This work is unique because it allows insight into how the full-coverage discrete film cooling layer is interrupted by high momentum dilution jets and how the surface cooling is affected. The film cooling flow was disrupted along the combustor walls in the vicinity of the high momentum dilution jets and the surface cooling effectiveness was reduced with increased dilution jet momentum. This was due to the secondary flows that were intensified by the increased jet momentum. High turbulence levels were generated at the dilution jet shear layer resulting in efficient mixing. The film cooling flow field was affected by the freestream turbulence and complex flow fields created by the combined dilution and effusion cooling flows both in the near dilution jet region as well as downstream of the jets. Effusion cooling holes inclined at 20Ë created lower coolant layer turbulence levels and higher surface cooling effectiveness than 30Ë cooling holes. Results showed an insensitivity of the coolant penetration height to the diameter and angle of the cooling hole in the region downstream of the dilution mixing jets. When high momentum dilution jets were injected into crossflow, a localized region in the flow of high vorticity and high streamwise velocity was created. When film cooling air was injected the inlet flow field and the dilution jet wake were fundamentally changed and the vortex diminished significantly. The temperature field downstream of the dilution jet showed evidence of a hot region which was moderated appreciably by film cooling flow. Differences in the temperature fields were nominal compared to the large mass flow increase of the coolant. A study of streamwise oriented effusion film cooling flow without dilution injection revealed unique and scaleable velocity profiles created by the closely spaced effusion holes. The effusion cooling considered in these tests resulted in streamwise velocity and turbulence level profiles that scaled well with blowing ratio which is a finding that allows the profile shape and magnitude to be readily determined at these test conditions. Results from a study of compound angle effusion cooling injection showed significant differences between the flow field created with and without crossflow. It was found from the angle of the flow field velocity vectors that the cooling film layer grew nearly linearly in the streamwise direction. The absence of crossflow resulted in higher turbulence levels because there was a larger shear stress due to a larger velocity difference between the coolant and crossflow. The penetration height of the coolant was relatively independent of the film cooling momentum flux ratio for both streamwise oriented and compound angle cooling jets. / Ph. D.

Higher turbine inlet temperatures (TIT) are required for higher overall efficiency of gas turbine engines. Due to the constant push towards achieving high TIT, the heat load on high pressure turbine components has been increasing with time. Gas turbine airfoils are equipped with several sophisticated cooling technologies which protect them from harsh external environment and increase their operating life and reduce the maintenance cost. The turbine airfoils are coated with thermal barrier coatings (TBCs) and the external surface is protected by film cooling. The internals of gas turbine blades are cooled by relatively colder air bled off from the compressor discharge. Gas turbine internals can be divided into three broad segments – Leading edge section, (2) mid-chord section and (3) trailing edge section. The leading edge of the airfoil is subjected to extreme heat loads due to hot main gas stagnation and high turbulence intensity of the combustor exit gases. The leading edge is typically cooled by jet impingement which cross-over the rib turbulators in the feed chamber. The mid-chord section of the turbine airfoils have serpentine passages connected via. 180° bends, and they feature turbulence promotors which enhance the heat exchange rates between the coolant and the internal walls of the airfoil. The trailing edge section is typically cooled by array of pin fins. On one hand, the coolant routed through the internal passages of turbine airfoil help maintain the airfoil temperatures within safe limits of operation, the cooled air comes at a cost of loss of high pressure air from the compressor section. The aim of this study is to develop internal cooling concepts which have high thermal hydraulic performance, i.e. to gain high levels heat transfer enhancement due to cooling concepts at lower pumping power requirements. Experimental and numerical studies have been carried out and new rib turbulator designs such as Criss-Cross pattern, compound channels featuring uniquely organized ribs and dimples, novel jet impingement hole shapes have been developed which have high thermal-hydraulic performance. Further, gas turbine blades rotate at high rotational speeds. The internal flow routed thought the serpentine passages are subjected to Coriolis and centrifugal buoyancy forces. The combined effects of these forces results in enhancement and reduction in heat transfer on the pressure side and suction side internal walls. This leads to non-uniformity in the heat transfer enhancement which leads to non-uniform cooling and increase in the sites of high and low internal wall temperatures. Development of cooling concepts which have high thermal hydraulic performance under non-rotating conditions is important, however, under rotation, the heat transfer characteristics of the internal passages is significantly different in an unfavorable way. So the aim of the turbine cooling research is to have concepts which provide highly efficient and uniform cooling. The negative effects of rotation has been addressed in this study and new orientation of two-pass cooling channels has been presented which utilizes the rotational energy in favor of heat transfer enhancement on both pressure and suction side internal walls. Present study has led to several new cooling concepts which are efficient under both stationary and rotating conditions. / Ph. D.

This paper investigates the effect of blowing ratio on the film cooling performance of a showerhead film cooled first stage turbine blade. The blade was instrumented with double-sided thin film heat flux gages to experimentally characterize the Nusselt number and film cooling effectiveness distributions over the surface of the blade. The blade was arranged in a two-dimensional, linear cascade within a transonic, blowdown type wind tunnel. The wind tunnel freestream conditions were varied over two exit Mach numbers, Me=0.78 and Me=1.01, with an inlet freestream turbulence intensity of 12% , with an integral length scale normalized by blade chord of 0.26 generated by a passive, mesh turbulence grid. The coolant conditions were varied by changing the ratio of coolant to freestream mass flux, blowing ratio, over three values, BR=0.60, 1.0, and 1.5 while keeping a density ratio of 1.7. Experimental results show that ingestion of freestream flow into the showerhead cooling plenum can occur below a blowing ratio of 0.6. Film cooling increases Nusselt number over the uncooled case and increasing the blowing ratio also increases Nusselt number. At a blowing ratio of 1.5 and Me=1.01 a large drop in effectiveness just downstream of injection on both the pressure and suction surfaces is evidence of jet liftoff. The blowing ratio of 1.0 was found to have superior heat load reduction over the blade surface at both freestream conditions tested. The blowing ratio of 1.0 reduced the heat load by as much as 39% and 32% at Me=0.78 and 1.01, respectively. / Master of Science