By Bjorn Fehrm

August 16, 2024, ©. Leeham News: We do an article series about engine development. The aim is to understand why engine development now has longer timelines than airframe development and carries larger risks of product maturity problems.

To understand why engine development has become a challenging task, we need to understand engine fundamentals and the technologies used for these fundamentals.

We have covered the different areas of a gas turbine except the exhaust (Figure 1). Before we go to the exhaust and the different outputs from a gas turbine Core, we look at the temperatures and how to cope with them in the different sections of a gas turbine.

Figure 1. The gas turbine cycle and its parts. Source: Rolls-Royce: The Jet Engine.

The Core Temperatures

The engine in Figure 1 is one of the early jet engines with moderate pressures and temperatures. We use it as an orientation graph, giving the names of the different parts of a gas turbine and how temperatures, pressures, and air speeds change when it passes a gas turbine. Its pressures and temperatures represent a jet engine from the 1950s.

To understand the temperatures and pressures of a modern engine, we use a recent GasTurb simulation I have made of the CFM LEAP-1A32, which is used in the Airbus A321.

Figure 2. The GasTurb graph of a twin-shaft turbofan of the LEAP type. Source: GasTurb.

Figure 2 shows the engine in a graphic representation, as used in GasTurb, to identify parts, cooling air flows, and the stations in the engine. Stations are a standardized way to identify the different parts of a gas turbine.

The engine area before the first compressor/fan stage is always called Station 2, the end of the compressor Station 3, the end of the combustor Station 4, the entry to the low-pressure turbine Station 45, and the end of the turbines Station 5. The core exit is Station 6, and the area where the bypass and core flows meet is Station 8.

If we look at the temperatures, pressures, and flow speeds that the different sections of the engine are subject to at a hot day’s takeoff (at the so-called corner point of the engine, the highest Outside Air Temp (OAT)), where the engine still gives maximum takeoff thrust), we get the diagram in Figure 3.

Figure 3. Temperatures and pressures at Takeoff for the LEAP-1A32. Source: GasTurb simulation by Leeham Co.

The values are the values at the Stations. For instance, the Mach number at Station 4 is 0.3, and at Station 44 (after the HPT) is 0.5. But this is the end value when the impulse+reaction turbine has extracted the pressure by diffusion of the airspeed from the HPT Nozzle.

We have a Combustor end – HPT Nozzle entry temperature (Station 4) of 3,800R (Rankine, the Imperial absolute temperature equivalent to Kelvin. I’ve added °C to the graph as well) or 2,110K/1,838°C and a static pressure of 575 PSI/3,964kPa. As the combustion gas has passed the LEAP two-stage High-Pressure Turbine (HPT), it is at 2,300R/1,278K and 100PSI/689kPa.

This means the high-pressure compressor, over the HPT shaft, has consumed energy equivalent to 475 PSI out of a total of 575 PSI of the mass flow of 85lb/s. The HPT absorbs 400 PSI at temperatures as high as 1,800°C, while the turbine blades and discs are subject to the centrifugal forces from a shaft maximum speed of 19,400 RPM (from the LEAP-1A FAA/EASA certification data sheet).

At the same time, the best Superalloy for gas turbine use, Rene 41, keeps its high strength to only around 700°C, Figure 4.

Figure 4. GasTurb materials database extract for Rene 41 Superalloy (the thicker top curve). Compare its strength versus temperature with Ti-6Al-4V (the typical Titanium alloy), the lowest curve. Source: GasTurb.

How to make the alloy survive an atmosphere 1100°C hotter than its Ultimate strength temperature with a gas speed of around 2,100 mph at the HPT Nozzle entry? It’s the subject of the next Corner.