Innovating turbine cooling

The START Lab evaluates advanced turbine cooling designs and methods

Improving Turbine Efficiencies through Heat Transfer and Aerodynamic
Research in the Steady Thermal Aero Research Turbine (START)

Department of Energy

The overall goal of this project is to advance cooling of gas turbine components with the aim of improving efficiencies and lowering costs. The specific goals for the project are to demonstrate increased turbine efficiency by reducing cooling flow to the turbine through the systematic studies of Reynolds number, cooling flow rates, and airfoil cooling designs and to determine the appropriate scaling parameters for different testing environments. The scope of the project includes: the design and manufacturing of a rainbow set of blades with baseline and advanced cooling technologies available in the public literature; measurements of aerodynamics and heat transfer for baseline and advanced configurations; continual assessment of additive manufactured components to reduce costs and advance cooling designs; and the development of the National Experimental Turbine.

As designers aim to increase efficiency in gas turbines for aircraft propulsion and power generation, spatially-resolved experimental measurements are needed to validate computational models and compare improvement gains of new cooling designs. Infrared (IR) thermography is one such method for obtaining spatially-resolved temperature measurements. As technological advances in thermal detectors enable faster integration times, surface temperature measurements of rotating turbine blades become possible to capture including the smallest features. Part of this project is to develop instrumentation including the use of the latest IR detector technologies for capturing spatially-resolved rotating blade temperatures.

A rainbow set of turbine blades will be cast with baseline and advanced cooling technologies available from the public literature for testing.

Thermal image of a START rotating cooled turbine blade with a moderate blowing.

The National Experimental Turbine (NExT)

Department of Energy

The goals for the National Experimental Turbine (NExT) are to develop an advanced turbine geometry for US companies and institutions that provides a platform for acquiring detailed data for design tool development. Once developed, the advanced turbine can be used by the industry partners to develop proprietary cooling designs, secondary seals, etc. that can be quickly and cheaply tested. Detailed baseline test data will be provided to the industry partners (Honeywell, PW, Siemens, and Solar Turbines). The majority of the funding for the turbine development is made available by DOE while a 20% cost share is provided by the industry partners.

The National Experimental Turbine (NExT) will be placed in the START Lab and will provide the industry partners with a test bed for evaluating novel cooling designs.

Leading Advanced Turbine Research for Hybrid Electric Propulsion Systems

NASA-University Leadership Initiative

Hybridization of aircraft propulsion, which is particularly well-suited for single-aisle medium- and short-haul aircraft, offers opportunistic benefits in reducing energy, carbon emissions, noise, maintenance cost, and cash operating cost. NASA has laid the foundation for a successful progression of hybrid electric propulsion systems that highlights the need to meet the challenges of smaller and highly efficient cores. The objective of the proposed work is to optimize the gas turbine design for hybrid electric propulsion to meet turbine efficiencies equal to that of large cores, and reduce the energy consumption of the turbine by 8-12% over the entire flight envelope.

ASCENT Project 56: Reduced Fuel Burn through Double-Wall Cooling of Turbine Airfoils Made Possible through Additive

Federal Aviation Administration

Gains in cooling performance of cooled turbine airfoils have a direct impact on the efficiency of turbine engines and therefore is the subject of much development. Today, many cooling designs for turbine airfoils use complex micro-channels placed in the wall of the airfoil to extract heat, which is otherwise known as double-wall cooling. The geometric complexities of the micro-channels, however, are limited by the casting process for these relatively small features. One approach to manufacture these complex channels is the use of metal-based additive manufacturing (AM), which is also known as laser powdered bed fusion. AM has begun to see many uses in the gas turbine industry, particularly because of the new design space enabled by this new fabrication method. However, not only does AM open up new design opportunities, but it also provides a quick turnaround from design concept to manufactured component relative to the traditional casting process.

The impact of this work is twofold: i) to develop an optimized micro-channel (double-wall) cooling design that will result in reduced cooling air but maintain turbine airfoil durability; and ii) to assess the viability of using additive manufacturing to print complex double-wall cooling designs. While optimization methods have been employed to develop new cooling technologies, the success of these technologies is only made possible if the designs can be made feasible. Thole’s previous research, has shown that it is feasible to use optimization methods to develop unique cooling features and although this previous research was not done for airfoil cooling, a similar methodology could be employed for this project.

CT scan of an optimized wavy channel that was built using AM.

AM can save significant money and time when making turbine blades.

Impact of Ceramic Matrix Composite Topology on Friction Factor and Heat Transfer

Pratt & Whitney

Ceramic matrix composites (CMCs) are of interest for hot section components of gas turbine engines due to their low weight and favorable thermal properties. To implement this advanced composite in a gas turbine engine, characterizing the influence of CMC’s surface topology on heat transfer and cooling performance is critical. However, very few published studies have reported the flow and heat transfer effects caused by this unique surface topology. This study is an experimental and computational investigation to evaluate the effect of weave orientations, relevant to CMC surfaces, on the resulting pressure loss and convective heat transfer within an internal channel. The weave pattern was additively manufactured as the walls of a scaled-up coupon containing a single channel. For each of the three weave orientations, bulk pressure losses and convective heat transfer coefficients were measured over a range of Reynolds numbers.

Illustration of a generic weave pattern for CMC studies (left) and predicted local Nusselt number contours for Re = 40,000 on the full length of the one weave wall coupons for a 0° orientation, 45° orientation, and 90° orientation (right).

Impact of Ceramic Matrix Composite Topology on Friction Factor and Heat Transfer

Pratt & Whitney

One of the most critical issues related to the operation of a gas turbine in today’s world is the ingestion of dirt and other fine particles that lead to blockages of cooling holes and passages required for effectively cooling the walls of the combustion chamber. Because the need to fly in dirty environments is on the rise, the criticality of operations in dirty environments is increasing. Modern gas turbine engines typically employ a double-walled combustor liner with impingement and effusion cooling plates whereby impingement cooling enhances the backside internal cooling and effusion cooling creates a protective film of coolant along the external liner walls. Dirt accumulation on the internal surfaces severely diminish the heat transfer capability of these cooling designs. Additionally, dirt accumulates on the external surfaces of combustor panels.

The impact of this work is twofold: i) to develop a cooling design for combustor liners that is insensitive to “dirty” cooling air; and ii) to characterize the cooling performance and the impact that the cooling design has on the downstream turbine.