Discussion

The supersonic flight problem

The transport of passengers in high supersonic speed is quite different from the normal cruise at around Mach 0.8. Concorde cruised at 2.0 and demonstrated to the world the kind of technologies that would be needed.

One of the areas which is very different to a normal airliner is the engine and its enclosure, the nacelle. A jet engine get its thrust by giving the air that it has swallowed a higher speed out the back (the air overspeed as we call it). Thrust is the sum of the mass of air being ejected times its overspeed.

At a normal airliner speed of M0.8, the aircraft flies around 300m/s faster than the surrounding air (M0.8 equals ~300m/s at FL300 or higher). As the engine’s compressor needs to work with air at an axial speed of ~M0.5 and the air entering the combustor should go even slower (ca M0.3), the incoming air is slowed down in the nacelle intake.

As air is slowed down, the air pressure increases, for a subsonic airliner to approximately 1.5 times higher than at the nacelle inlet. As air pressure increases, the air temperature increases; for the normal airliner with 30°C.

A supersonic airliner flies at M2.0-2.2 (Concorde at M2.0, Boom at M2.2). Here, the air needs to brake from 600-660m/s to around 150m/s at the engine inlet. Should this slowing of the air happen without losses, we would have a pressure rise of 10.7 times. Unfortunately, we also have an air temperature increase of 170°C.

Materials for making compressor parts/casings can withstand around 700°C. For a normal airliner engine, the -56° outside air temperature at FL360 (11,800m) and higher and the 30°C rise before the compressor leaves 726°C for the compressor temp. increase and therefore air compression. This allows compression ratios of 110 times if we assume the compressor works with the common 90% (isentropic) efficiency.

The most advanced long-range engines today use a maximum compression of 58 times (GE GEnx-1B75), so there is plenty of margin before we hit a compressor temperature limit at cruise.

This is not the case for the SST engines. In fact, the compressor temperature and pressure at the nacelle intake influences the whole design of the engine. We must keep compression levels down in the engine to make things work.

Boom SST engine

Boom Technology stated that it identified existing engine cores that could work for their SST engine. This engine core should be combined with a new low pressure system, i.e., fan/low pressure compressor/turbine to be paired with the core’s high pressure compressor/combustor/turbine.

The problem is that we can only allow the new combined system to raise the pressure by 23 times, otherwise we exceed the material limit of 700°C at the end of the compressor.  We only find 23 times Overall Pressure Ratio (OPR) in supersonic fighter engines for the same reason as for the SST.

The Eurofighter Typhoon has an engine, the EJ200 (Figure 2), which has a suitable compression ratio of 26 in total; four are in the triple fan section. It also has a core that could be suitable for our SST engine. The core’s high pressure compressor has a pressure ratio of 6.5.

Figure 2. Eurofighter EJ200 turbofan with afterburner and advanced nozzle. Source: Rolls-Royce.

Inlet

The Typhoon has the same multi-chock inlet type that the SST need to fly efficiently at M2.0, giving the intake system a pressure gain around 10 (and a temp. increase of 170°C). The overall pressure ratio (OPR) of the engine is adapted to this flight situation.

The Typhoon and the Boom SST need an advanced inlet for the nacelle to gain the incoming air’s latent pressure ratio of over 10 at M2.0. Figure 3 shows what happens if we use a conventional pitot intake like on present airliner engines.

Figure 3. Intake losses for multi chock inlets like F15/Concorde/Typhoon and std. pitot intake like F16 or jet airliners. Source: Cumpsty “Jet Propulsion”.

Like the F16, we would lose pressure but not temperature as energy from losses converts into increased temperature of the air. A pressure loss of 20% at M2.0 compared to a variable multi-chock inlet (like the F15) is not acceptable. Hence, our SST nacelle must have a variable multi-chock inlet, like the F15/Concorde, Figure 4.

Figure 4. A variable ramp multi-chock inlet like the Concorde/F15 inlet. Source: Cumpsty “Jet propulsion”.

Mixer and nozzle

The EJ200 engine in Figure 3 also has something else which is needed on a SST engine: an advanced Convergent-Divergent outlet nozzle. At an aircraft speed of 600m/s, we need to accelerate the air leaving the engine at a speed higher than 600m/s, otherwise we will not get any thrust.

The air inside the engine has a speed of around 150m/s and we need to accelerate the air to at least 700m/s to get any overspeed value in our overspeed*mass_of_air thrust equation. A conventional fixed Convergent nozzle is a compromise between efficient speed increase of the air and low weight/complexity.

A Convergent-Divergent nozzle works much better, to the level that it’s necessary for a SST application. To accelerate all the air, both core and by pass stream, we need to mix the air before the Convergent-Divergent nozzle in a mixer. For the EJ200, the mixer also doubles as an afterburner, something that the Boom SST engine will not use.

SST engine design and data

In the next article, we will dig deeper into the SST engine. We will examine the data for the EJ200 core and see if it could be used for our SST engine. If so, what changes would need to be done?

We will also consider an alternative engine core, the GE414, and see if this would suit our needs better.

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