Bjorn’s Corner: Supersonic transport revival, Part 10

October 12, 2018, ©. Leeham News: In the last Corner we discussed the challenges of the nacelle outlet for an SST (SuperSonic Transport). Now we will discuss SST engines and what are the key technical challenges for these engines.

We start this week by looking at some design constraints for the engine which we don’t have in Subsonic airliner engines.

Figure 1. A generic mixed Turbofan SST engine with ConDi nozzle. Source: GasTurb.

The key design challenges for an SST engine

Some Corners ago we discussed the challenge of Ram drag for an SST engine. It will limit the Bypass ratio we can have for a fix cycle engine (engines which don’t have a variable Bypass ratio for take-off and cruise, called variable cycle engines). We will come back to the Ram drag and Bypass ratios when we discuss noise in a later Corner. Now we look at some other challenges in an SST engine.

The compressor exit temperature challenge

For Subsonic airliner engines, the key hot area is the first turbine stage after the combustor.  In Figure 1 it’s the T41 area, the Turbine Entry Temperature (TET) point. The higher this temperature is, the more power the core produces per kg air which passes the core. And the more powerful the core, the larger fan it can drive over the turbines.

The ideal is a small core which drives a large fan to give a high bypass ratio, which gives a high propulsive efficiency. At a speed of below M0.85, the resulting Ram drag is not a problem. This changes drastically for SST engines as we will see in later Corners. Now we focus what happens in the engine’s hot spots.

The limit of T41 for a small and efficient core is set by available materials but also how we can use cooling air in and around the turbine blades and stators. Suitable cooling air with a high enough pressure to feed the turbine section is available from the last stages of the compressor.

An SST engine can use the available turbine technology developed for the Subsonic airliner engines, the turbine section is therefore known technology and presents no special problems.

For airliner engines flying at M0.85 or less, the end temperature of the compressor (T3 in Figure 1) is not a problem. This is good because there is no cooling air available (the last stage of the compressor, which is the hottest, has also the highest pressure level in the engine, so no other air can cool it). The available materials (nickel-based alloys) can handle the temperatures of the last compressor stages without cooling.

For an SST engine, the maximum allowed T3 temperature is a problem. The faster we fly the higher the temperature at the entry of the engine and the higher the T3 will be with a certain level of compression. At high Mach, the speed decides the maximum compression ratio the engine can have.

The maximum temperature for the final stages of a compressor is around 700°C. With the air entering the engine at 115°C after our multi-chock inlet at Mach 2.2 ( compared with -25°C for a subsonic M0.85 airliner), we can have a total fan+compressor Overall Pressure Ratio (OPR) of 23 (compared with over 100 for an airliner engine).

With the fan section delivering around PR 3.5 (it must be at this level to have the bypass air match the core’s exit pressure in the mixing zone at Station 64 in Figure 1) we search for a core with a high-pressure compression of less than 7. Such airliner engines were designed in the 1970ies.

Today we have high-pressure compressors at more than PR 15, yesterday at 10 (the CFM56 for example). Cores with a high-pressure compressor at less than 7 is only found in military engines which are designed for flying at Mach 2 or above. The conclusion is existing cores of the right size suitable for use in a Mach 2.2 SST are very hard to find.

If we instead fly at Mach 1.4 we can have a pressure ratio of the high-pressure compressor of 10. The CFM56 core can then be used, which is what GE did for the Aerion SST engine. A suitable low-pressure system was designed for this core with a fan section pressure ratio of about 2.8.


We have now looked at the critical areas Ram drag, compressor exit temperature and allowed pressure ratios for high-speed SSTs and what this means for compressor designs and suitable existing cores.

Next week we will look at the noise problem which is intimately coupled to the Bypass ratio and therefore Ram drag.

16 Comments on “Bjorn’s Corner: Supersonic transport revival, Part 10

  1. Björn, to my Knowledge it is no longer true that modern subsonic turbofans are not T3 limited, for the reasons you quote (high temp, no possibility for pressure cooling, difficult to use non-metallic materials). Maybe an Engine expert could comment?

      • Thanks Dukeofurl, So the next generation engines from GE raises the T3 bar from 700°C to 760°C. It has a PR 27 high-pressure compressor and an OPR of 64, which is the reason GE found new materials for the last stages to raise the bar to 760°C.

        The problem for the SST projects is they are looking for PR 10 (Mach 1.4) or even a PR 6 core (M2.2) and these were not made with 760°C T3 materials. One can be lucky if the design handles 700°C.

    • Hi James,

      how nice, a Concorde engine expert. Please chip in with your experiences! What was the highest T3 on that engine?

  2. Hi Karle,

    I have pretty good GarTurb models of all the modern engines, and you are right. The highest pressure ratio ones (specifically the GEnx-1B76) push the T3 limit of 700°C in the rare case of a One Engine Inoperative full thrust situation at V2 Safety Speed and ISA +15 temps. In cruise, they are below 550°C and at Top Of Climb below 650°C. So for everyday use, only the SST engines are having a real problem with T3.

  3. As a historical anecdote, the Spey powered Phantom, while superior to the J79 version in all respects except max speed at altitude. This was because the civilian derived Spey low bypass turbofan was limited by compressor outlet temp under those conditions.( According to Flying Review International magazine from Aug 1966). The components could have been redesigned but at extra cost and time delay.
    Seems to reflect the same issues Bjorn has covered when a civil engine is adapted for supersonic use

      • It was superior as the engine provided more dry thrust, and as being a turbofan the AB gave a higher maximum thrust. These numbers are uncontested and the other numbers were:
        15% increase ferry range
        10% increase in combat radius
        30% decrease in takeoff distance (runway)
        climbing time to 40,000ft reduced by 1/3
        Max speed at altitude came down to around M 2.0 for reasons as discussed.
        A longer extending nose-wheel increased performance off the catapult.
        Of course the rest of the plane was essentially the F-4J version, so had a very good platform to go from. The USAF seriously looked at the Spey version for its fighter-bomber variants but decided against
        it but a developed non AB version of the Spey as the Allison TF41 replaced the problematic PW TF30 on Voughts A-7.

        A similar improved performance happened with the F16 when the early engine was replaced by higher powered PW F100 and GE F110 engines.
        Same applies to civil aircraft, more power in a newer design generally means better performance. Works too for turbo props like the Hercules or business jets. It goes without saying

        • compressor outlet temps:
          How was this handled on the Saab Viggen
          using the civil airliner JT8D based Volvo RM8 ?

          • The JT8 was modified by Volvo Flygmotor into RM8. It had an OPR of only 16.5, meaning the Temp increase through the fan/booster/compressor with a Mach 2 inlet temp of 116°C (you could only fly at Mach 2 above 39,000ft where the outside temp is -57°C) was 610°C, below the 700°C capability of Inconel.

            BTW, GE gave more info around the AS2 engine, called GE Affinity, at the NBAA biz jet show in Orlando today. It has two fan stages for the higher bypass pressure I told about. But the high-pressure compressor has cut one stage from the CFM56 (from 10 to 9) to lower the compression ratio, probably the first stage and probably because of T3 concerns (Edit the week after: My memory tripped me, the CFM56 has 9 HPC stages so no change for the Affinity). The engine looks much like the sketch in Figure 1 but with a fan with 2 stages and a fixed nozzle. Doable at Mach 1.4 as discussed in Part 9. We will look more into the Affinity in coming Corners.

  4. Wonder if the USAF would be interested in the AS2 ,for some missions, powered by an engine with these sort of features

    ‘Features include a service ceiling of 60,000 feet, the highest bypass ratio of any supersonic engine, a proprietary non-augmented exhaust system (no afterburner required), a combustor with advanced coatings optimized for sustained high-speed operations, advanced acoustic technology to meet or exceed regulatory requirements, and additive manufacturing technologies to reduce weight and increase performance.’

    • It gets to be tricky as normally the whole design and its Engines has to be USAF certified according to their specs with damage tolerance etc… Like the KC-46A derived from the 767.

  5. GE normally cuts the last stages of the HPC when they want to increase core flow and reduce T3 temp. The CF6-6 thru CF6-80E1 HPC is a good study. To handle the high T3 temperatures that go buyond Inco718 capabilities you get into forged powder metals, Udimet 720 was one and today there are lots of better versions pushing T3 over 1200F for high stress applications.

      • The TF-39 had 16-stage HPC
        The CF6-6 had 16-stage HPC, then came the switch with 2 less HPC satges to increase core flow and adding stages on the booster. You get more Power but loose bypass ratio.
        The CF6-50 had 14-stage HPC
        The CF6-80A had 14 stage HPC
        The CF6-80C2 has 14 stage HPC
        The CF6-80E1 has 14 stage HPC
        GE90-94B has 10 stage HPC, same trick here, to remove last HPC stage to increase core flow.
        GE90-115B has 9 stage HPC
        I can recommend GEAE book 8 decades of progress.

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