, a good question!
Occasionally I write something, which I either know or believe to be correct, but which, as I write it, I rather hope I am not asked to justify.
When I wrote…Mach 3 would have been ideal from an engine point of view…
that was one of those occasions, and I also have to admit, in retrospect, I did not make it clear that I was talking theoretically.
I will do my best to answer, but let me say at the outset that I am not an expert on the RR
Snecma Olympus 593 Mk 610 engine, and so I have consulted BAe and RR
technical notes for information.
It would have been better, and probably more accurate, had I said …Mach 3 would have been ideal thermodynamically from a theoretical engine point of view…
The ideal engine for cruise at moderate supersonics speeds is a relatively low pressure ratio pure turbo jet engine. Pure jet because we need the engine to be capable of being installed in a slender nacelle, and relatively low pressure ratio because the intake itself will make a major contribution to total compression, about 7:1 at Mach 2.
At Mach 2, due to the design of the intake and exhaust systems, the core engine need only develop about 50% of the total thrust required, as the intake system and the exhaust system both contribute 25% each. This allows Concorde to fly with the re-heats off at Mach 2, substantially reducing fuel flows and increasing range.
However, this simple picture is complicated by two main points. Firstly, such an engine, designed and optimised for supersonic cruise, will probably not be powerful enough for take-off or transonic acceleration, and secondly, fuel consumption, whilst flying away from the design cruise regime cannot be ignored.
The first point can be solved by using re-heats as a method of augmenting basic engine thrust levels efficiently, with the advantage that it does not increase nacelle frontal area and has very low internal losses.
The second point however leads to a compromise being reached, and a higher pressure ratio engine being used than would be suggested by Mach 2 cruise considerations alone.
If we now take this engine to Mach 3, this total compression ratio rises to about 34:1, in theory a substantial achievement, which should result in even less thrust being required from the core engine, leading to my original comment.
However there would be problems in other areas, principally that while this very high pressure ratio is, in theory, desirable thermodynamically, the temperature rise taking place during compression, combined with the physical properties of the materials used in the compressor blades, would limit the amount of fuel which could be burned, which would in turn limit the specific thrust.
Put simply, the compressor blades wouldn’t have handled the temperatures, let alone the turbine blades.
So, as temperature was causing major problems in other areas of the aircraft design, such as cabin air cooling, fuel, oil and hydraulic fluid temperatures, and the design Mach number kept coming down, eventually to Mach 2.04, the engine designers saw no need to solve the problems of Mach 3, as the airframe was never going to get there.
…Or did you mean that it would have been the maximum attainable speed with the engine technology available?...
Even at Mach 2, the latter stages of the compressor are made of a nickel-based alloy, normally reserved for the turbine area, and though the engines could have handled the temperatures at speeds a little above Mach 2, it seems unlikely that Mach 3 could have been achieved at that time, with the materials then available, however desirable it may have been theoretically.
I’m sorry if I confused you initially, and I hope that this explanation has not made matters even worse.
The following sources gratefully acknowledged:
Chris Orlebar “The Concorde Story”
British Aerospace PLC Training Centre “Concorde Conversion Lecture Notes”