The new Cooper-Bessemer power turbine is a high-efficiency, center frame-mounted, three-stage unit that can be driven by either the existing RB211-24 gas generator or the new improved version. The upgraded gas generator combined with the new power turbine offers an increase in nominal output from 28.4 MW (38,000 hp) to 31.8 MW (42,600 hp). The new coupled turbine, now being tested, is called the Coberra 6761. Besides improving core engine performance, the program's objectives included improved fuel efficiency and reliability, and easier site serviceability; extension of the modular concept from the gas generator into the power turbine with improvements in sealing, materials, and temperature capability as well as interchangeability of both upgraded turbines with existing hardware. The Rolls-Royce industrial RB211 turbine, derived from an aircraft engine, is the basis for the gas generator end of Cooper Energy Services' Coberra coupled turbines. The power turbine design capacity has a significant effect on the power at a given speed. The flow capacity was optimized to achieve the best thermal efficiency and lower IP speeds to optimize IP compressor efficiency and permit future throttle push.
Transmitting supplies of natural gas over long distances through pipelines requires pumping stations every so often along the route, to make sure that the gas continues to flow.
Currently, the most widely accepted approach to generating power for these applications involves using gas turbines derived from those used in aircraft, powered by the very gas being pumped. A common arrangement involves coupling such a turbine, operating as a gas generator, with a power turbine to drive the gas compressor. This is the approach followed by Cooper Energy Services of Mount Vernon, Ohio, a division of Cooper Cameron Corp. of Houston, with its Coberra 6000 range of coupled turbines, which incorporate the Rolls-Royce RB211-24G aircraft-derivative gas turbine and a Cooper-Bessemer power turbine from Cooper Energy Services.
The economics of new projects today favor ever-larger power blocks and higher thermal efficiencies. Accordingly, it was thought advisable to further develop the Coberra system. This involved not only upgrading the gas generator by the selective introduction of technology from RollsRoyce's Trent engine, but also the introduction of a new power turbine from Cooper-Bessemer. The Trent aircraft engine has approximately 50 percent more thrust than the highest-rated RB211 aircraft engine, although it maintains the older engine's configuration and much of its engine geometry, with the result that many of the Trent's more efficient components can be used in upgraded variants of the RB211. The new Cooper-Bessemer power turbine is a high-efficiency, center frame-mounted, three-stage unit that can be driven by either the existing RB211-24 gas generator or the new improved version. The upgraded gas generator combined with the new power turbine offers an increase in nominal output from 28.4 MW (38,000 hp) to 31.8 MW (42,600 hp).
The new coupled turbine, now being tested, is called the Coberra 6761. Besides improving core engine performance, the program's objectives included improved fuel efficiency and reliability, and easier site serviceability; extension of the modular concept from the gas generator into the power turbine with improvements in sealing, materials, and temperature capability; and interchangeability of both of the upgraded turbines with existing hardware.
In seeking the most effective upgrade route, the engineers on the project were aware that power can be increased either by increasing the firing temperature or by increasing the flow through the engine. But increasing the flow usually means significant component change, either to increase the area of the compressor intake annulus or to add a "zero" stage in front of the existing stage 1 rotor. On the other hand, raising the firing temperature by a large amount can lead to an increase in undesirable emissions. Adding a zero stage was rejected, since the additional length of the rotor and casing path changes the engine dynamics, and this would increase the technical risk and compromise interchangeability. This option also would increase the pressure and temperature at the entry to the combustion chamber, and would require the redesign of the following stages to optimize the pressure characteristics. Fortunately, it was found that aerodynamic techniques recently developed by Rolls-Royce would permit the designers to modify the existing intermediate-pressure (IP) compressor to pass more flow with minimal mechanical change. This, along with a modest rise in firing temperature (well within aircraft engine experience), was enough to meet the performance requirements.
The IP compressor modification was intended to reduce the speed at a given level of flow, in order to increase efficiency and ensure the potential of the compressor to be further uprated. The intake case, outer casings, and disc pack are unchanged. The airfoils of the first two stages of the IP compressor have been redesigned in the style of the Trent aircraft engine. Tailored airfoils with controlled diffusion and reduced shock losses have allowed higher-duty stages 1 and 2 to produce the required flow increase with increased efficiency and adequate stability margins. Stage 3 and 4 vanes have been redesigned to accept the increased flow. The revised rotor airfoils are also lighter than the existing design, giving reduced serration loading. The outlet guide vane had previously had a high negative aerodynamic incidence. A reduction in this incidence, along with the introduction of a 3-D airfoil design, will allow better aerodynamic matching to the high-pressure (HP) compressor. The 3-D airfoil provides improved efficiency because it more fully addresses the boundary conditions, and since the resultant form is stiffer, thinner profiles can be adopted.
The HP compressor is to be replaced with an adaptation of the Trent 800 HP compressor, which offers a significant gain in efficiency, thanks to its advanced airfoils with controlled diffusion and revised work distribution, along with improved disk sealing and a low-windage drum construction. The drive arms and casing location did require minor reconfiguration to match interface locations.
The new engine is to be introduced initially as a gas-fueled dry low emissions (DLE) system, with non-DLE versions and dual fuel capability to come later. The combustor, the industrial RB211 DLE design, has nine radially mounted individual combustors, in an arrangement that maintains the existing gas generator length and rotor systems. The premix lean burn series staged system allows a significant operating envelope with controlled emissions. The combustor will be able to achieve 25 volume parts per million NOx and CO at the 24G upgrade design point.
The gas generator turbines of the old RB211 featured interlock shrouds, but to allow an increase in firing temperature without reduction in turbine life, the latest Trent interlocked rotor blade is incorporated in the new machine. These rotor blades feature an improved cooling configuration and a single crystal material that makes it possible to operate at increased firing temperatures. This HP rotor blade has already seen extensive use in aircraft engine settings.
The flow capacities of the IP and power turbines have been optimized. The IP turbine capacity is achieved by a small machine skew to the existing IP nozzle guide vane. The rotor blade is also now specified in single-crystal material to enhance life at increased temperatures. The HP/IP bearing housing has been changed from a two-panel design to a single-panel structure. This gives improved cooling air distribution around the bearing housing, leading to a lower operating temperature, a wider tolerance range, and an improved service routing.
Since upgrading the existing RT56 and RT62 power turbines would not have produced the gain in efficiency that was being sought, it was decided to introduce a new turbine using the latest Rolls-Royce aircraft engine low-pressure (LP) technology to maximize efficiency and mechanical integrity.
The first question to be considered in the redesign was what speed ranges were required for mechanical drive applications. Two types of oil and gas installations predominate: pipeline boosters that are directly driven and barrel compressors driven through speed-increasing gears. Since the optimum speed for the barrel compressor applications can be easily obtained by ratio changes in the speed increaser, it was thought best to select a power turbine design speed that maximizes the efficiency of existing pipeline booster frame sizes. Pipeline booster efficiency in the 30-plus MW (40,000-plus hp) power class is highest in the 4,500- to 5,500-rpm range.
The best power turbine efficiencies are obtained not only by moderate blade loading, but also by a low leaving loss associated with the blade path exit velocity or Mach number. Leaving loss is inversely proportional to annulus area, but rotating component stresses are directly proportional to annulus area and speed squared. Thus, for any combination of mass flow and desirable stress level, an optimum turbine design speed can be computed. It is also imperative to make the correct choice of number of stages; after considerable study, it was concluded that a close-coupled three-stage turbine at 4,850 rpm was the optin1llm solution.
Work split among the three stages was optimized for maximum overall turbine efficiency, with the first stage contributing the most and the last stage the least. As a result, the first-stage rotor blade metal temperature was reduced, and the last-stage blade path exit swirl entering the exhaust diffuser was minimized, thus helping to lower exhaust system losses.
The low aerodynamic loading of the three-stage turbine made it possible to achieve the goal of a close-coupled turbine, with the interturbine duct eliminated. The first-stage nozzle vane outer platform begins at the gas generator exit. The power turbine and gas generator are contrarotational, so gas deflection and the associated profile loss across the first stator are minimized. The resulting blade path produces a very high efficiency, yet retains the traditional robustness of industrial design.
The power turbine design capacity has a significant effect on the power at a given speed. The flow capacity was optimized to achieve the best thermal efficiency and lower IP speeds to optimize IP compressor efficiency and permit future throttle push.
To achieve its high efficiency, the power turbine incorporates new aircraft engine LP turbine aerodynamic technology and design techniques. One of the principal technology inputs involves "multilean" stators that use a blend of successfully demonstrated low-aspect-ratio airfoil design techniques with computational fluid dynamics evaluation within the aerodynamic design process. Another is curved airfoil stacking, an extension of the multilean concept to higher-aspect-ratio components operating in a sloping annulus. The technology also includes end-wall treatment to minimize the effects of airfoil passage secondary flow.
Also important is 3-D row-by-row flow matching, where the airfoil design is based on a multi stage coupled blade-to-blade streamline curvature through-flow model that includes airfoil passage secondary flow deviations in addition to the basic mainstream flow effects. Each component is defined within a fully 3-D viscous CFD model. This process enables simultaneous 3-D optimization of each airfoil and correct matching to adjacent ones.
Finally, there is the shape of the blade. The airfoil characteristics are similar to those on the Trent aircraft engine, and there is no separation at or near the suction surface throat point. The leading edge loading has been kept modest to guarantee good off-design performance, and all blade rows have been designed to accept substantial levels of incidence.
The power turbine design was organized into five easily separable modules-inlet duct module, turbine module, frame module, exhaust module, and bearings-to ensure control of concentricity during manufacture and to minimize maintenance downtime. This arrangement permits all critical or delicate assembly operations to be performed in the controlled conditions of a central service facility. Only the two hottest modules, the inlet duct (which includes first-stage stator vanes, the gas generator connection, and the inner diffuser wall) and the turbine module (main turbine case, second- and third-stage stator vanes, interstage diaphragms, and the entire rotor), will have to be exchanged at the site for routine maintenance, and their alignment with the rest of the machinery will not be disturbed.
The first-stage stator vane case is directly connected to the gas generator. This casing is machined from a rolled ring case. The main turbine case is a near-net-shape one-piece investment casting, so the number of flanges is reduced to the minimum needed for module separation. Extensive thermal analysis, applied to a finite-element model of the case, was used to design the casing liners to control operating temperatures of the load-carrying casings. Hollow stator vanes let cooling air into the rotor spaces.
To obtain closer clearance sealing at the disk rim and blade tips, a strutted support frame with seven struts was introduced to improve control of concentricity between the rotor and stator structures of the power turbine. The strut and frame design evolved from finite-element analyses for dynamics and thermal stresses, including steadystate and transient conditions using a fully 3-D model, and the structure is predicted to have an infinite life. The frame eliminates the need for an aft support of the bearing case and simplifies base construction, so the power turbine is actually lighter in weight than traditional models.
The airfoil design—stacked with axial and tangential leans to smooth the 3-D flow field-was also subjected to rigorous finite-element analysis. All three rotor blade rows have positive interlocking tip shrouds running against honeycomb-faced casing liners. The rotor blades are mounted on extended necks to reduce heat transfer into the discs. The three rotor discs are joined by precision-ground face couplings that transmit torque and maintain concentricity during transients; they are joined to the shaft by 12 through bolts.
The bearing cases are bolted to the strutted frame support with vertical joints sealed with O-rings, eliminating the need for horizontal joints or field-applied soft sealants inside the exhaust diffuser tunnel. Rotor bearings are of the traditional babbitt-faced hydrodynamic tilting pad type. All oil supply, vent, and drain services are connected on the cool aft end of the bearing case; no oil or air service lines cross the exhaust diffuser. A 1 ,800-rplTl auxiliary drive gearbox is provided for an oil pump drive.
Testing of the Coberra 6761 is now beginning, and the product is expected to be commercially available in about a year.
Air is shown streaming around one of the blades of the new Coberra power turbine in this CFD rendering.