I think you will find that because the engine's compressor can only handle incoming air at subsonic speeds ,then when the aircraft is at Mach 2.0 the 14 ft long intake has to slow the air down in rough figues from 1350 mph to
350 mph .
As the figure in your web site shows it does this by creating a number of shock waves in the forward part of the intake with the last shock wave[ where the air goes from supersonic to subsonic] being between the two intake ramps
In slowing the air down it also increases it's pressure , and I think the compression ratio was 7 to 1 but could be wrong. This was like having another compressor on the front of the engine with no need for turbine at the back
You can think of it like a piston engine with a turbo charger in the air intake so making the engine more powerful because it pre- compressed the air for the engine
Guys, another interesting Concorde thread (). OK, the thrust thing. YES, the numbers are just about correct, what we have is propulsive thrust that without a perfectly matched engine/intake would not be present. The divisions of the thrust are (Quoted from the publication "The Concorde Air Intake Control System").
The engine itself now only generates 8% of the total thrust, a mere shadow of its subsonic glory. The now divergent secondary nozzle produces a sizeable 29%, this being produced in a similar way to how the intake subsonic diffuser produces its thrust. (The main difference in the case of the secondary nozzle is that instead of a subsonic decelerating flow, we now have a supersonic accelerating flow). A huge 75% OF THE TOTAL THRUST is produced by the intake subsonic diffuser section, this being due to the huge rise in static pressure that is occurring in this section. The 'negative thrust' from the forward ramp section this time is 12%, produced by the supersonic compression forces acting on the divergent section of the intake, resulting in an intake thrust component of 63%. So it can be seen that the vast majority of the Mach 2 thrust forces are transmitted to the airframe not via the engine mountings, but via the mountings of the intake, and to a lesser extent the TRA nozzle. It might seem that the two cases, and in particular the latter one, are very demeaning to the role of the engine, but nothing could be further from the truth. By the laws of conservation of energy, thrust (or any other force for that matter) cannot be created out of thin air, the whole process is about maximising the powerplant thrust that is potentially 'on tap'. (O.K. I know, this entire subject is about providing thrust from thin air!!). Without the engine, the entire process of course falls apart and all components of the powerplant produce exactly the same amount of thrust - ZERO!! It is also doubtful if any engine currently in existence could do the supersonic job anywhere near as effectively as the OLYMPUS 593. (Not bad for a design that can be traced back over fifty-four years!). The 593 produces the necessary gas flows to produce these stated levels of thrust, and in the final analysis all powerplant thrust of course is really generated by the engine, what we have been looking at how this thrust is transmitted to the airframe.
I hope that this clarifies things guys, regards to all.
My previous post was written before I read yours and already hinted at the perpetual motion or free lunch situation that develops if the pressure rise-drop along the engine isn't accounted for correctly.
After all kerosene burning furnishes the power = drag x air speed.
See also: "Can Vmg exceed the V of a jet exhaust?"
There was a similar discussion last year, Can Vmg exceed the V of a jet exhaust?, in which ChristiaanJ (I think) reasonably criticized my dislike of the `free lunch'-type of description of supersonic intakes, owing to having been there and knowing the load-bearing structure of the concorde intakes.
The pressure of the air increases going backwards through the intake, and this produces a net forward force on the intake ramp. As M2dude's quote rightly states, the engine is enabling this to happen. Turn off the fuel and the thrust from all bits of the intake-engine-nozzle system no longer occurs.
Hang about - hang about!!! I'm sorry ..... do WHAT??
Look, you must forgive me - I'm not an aero engineer, just a humble MRI/CT scanner engineer and I am having very serious problems understanding this. Okay I understand perfectly that the 593, capable as it is, cannot accept supersonic air.
I further understand approximately that the intake is a prize winning work of engineering genius all on its own and that it is capable of leaching a whole 1000mph from the incoming air, so the 593 can do its thing. I also - just - understand that the aerodynamics of the nozzle does a huge amount of magic to the raw jet exhaust and appears to produce a thrust gain over and above the 593.
So far so good.
I was holding it all on the island because I reasoned that all that air going in has to go somewhere and if it was being slowed down, it must therefore be compressed - like another very powerful front end stage to the compressor. Then, I thought, that the compressor would be compressing compressed air so, although the engine gain would be the same, the result would be dramatically greater. Yes? Or no?
If not, by exactly which mechanism is real thrust being generated and, more importantly, transmitted to the airframe?
I love this aeroplane and since M2Dude, ChristiaanJ, Bellerophon, BS312, Exwok and all the other 'Concorde family' contributors have pitched in so generously, it is really coming alive. But she's damn complicated.
I hope that this diagram just might make matters a tiny bit clearer. It clearly shows how the propulsive thrust is divided up among the various components, particularly the intake. As the airlow travels through the carefully controlled and complex inlet shock system, it exhibits a 600% rise in Static Pressure Ps, this huge pressure rise reacts against the divergent wall of the intake, giving us colossal amounts of thrust. Take a look at this quote from ' The Concorde Air Intake Control System:
A good impression of the efficiency of any engine/intake combination can be gathered by looking at overall intake pressure recovery, as this will determine compressor face total pressure, itself being a major parameter in determining powerplant thrust. The following example is given for the A/C just before Top of descent, Mach 2.0, ISA +5 (This equating to temperatures of Ts = -51.5 ºC, Tt = 127 ºC) and altitude = 60,000'
Freestream Total Pressure = 8.14 P.S.I.A
Freestream Static Pressure = 1.04 P.S.I.A
Freestream Dynamic Pressure = 7.10 P.S.I.A.
Compressor Face Total Pressure = 7.63 P.S.I.A
Compressor Face Static Pressure = 6.42 P.S.I.A
Compressor Face Dynamic Pressure = 1.15 P.S.I.A.
Compressor Face Total Temperature = 127 deg's. C
Analysis of the above-described case shows that there is:
A SIXFOLD INCREASE IN STATIC PRESSURE !!!
AN INTAKE PRESSURE RECOVERY OF ALMOST 94%. THIS IS AN EXCEPTIONALLY HIGH FIGURE, PRODUCED WITH A STATIC PRESSURE INCREASE OF 5.38 P.S.I.
A REDUCTION IN DYNAMIC PRESSURE OF 6.78 P.S.I. NO HEAT ENERGY IS LOST IN THE COURSE OF THE COMPRESSION PROCESS
The pivotal part of all of this is a staggering 94% recovery of the freestream total pressure (the pressure coming at the intake), this is what enables the engine to move the required amount of airflow, enabling the intake, engine and nozzle to provide the thrust required for supersonic engine operation without the use of reheat and with staggeringly low fuel flow values.
This diagram shows just how the action of the two nozzles is able to help provide so much thrust at Mach 2. The 'cooling air' from the engine bay into the nozzle annulus s the secondary airflow that was diverted the intake ramps, and this gives the high pressure efflux an aerodynamic cushion to expand against from within the divergent secondary nozzle buckets. This gives a dramatic reduction in the thrust that would otherwise be wasted due to the high pressure efflux over-expanding/flaring against the very low static air pressure.
Guys, apologies if it seems that I'm trying to hog things here, I'm just attempting to help clarify an extremely complex and bewildering subject.
Just imagine for a second that we are flying (or attempting to fly) at Mach 2, but instead of having a convergent/divergent intake, we just have a hole at the front. (This is termed 'a pitot intake'). Contrary to common folklore supersonic air will not enter the engine, this is a fallacy. The velocity of air entering a jet engine compressor is defined by engine mass-flow demand and the cross sectional area of the L/P compressor, you cannot force air into a jet engine at a velocity it does not require. (You can certainly force the engine to surge, and possibly drive the inlet into unstart by attempting this though). The Olympus 593-610 at Mach 2 ISA +5 had a demanded compressor Mach number (Mn1) of 0.46, this is fixed. With just a pitot intake, what WILL happen is that to satisfy the demand of the engine a single normal shockwave will form across the face of the intake, resulting in subsonic flow downstream of the shock. (A normal shock will always without exception produce subsonic downstream air). Now the pressure losses involved with a normal shock are proportional to the 'strength' of that shock, where there is only a single normal shock is utilised, over 40% of the propulsive thrust would be lost at Mach 2. (Due to enormous compressor face distortion, the engine would also be unstable to the extreme. As Mach number increases this loss also increases, to the point that if we were able to fly at Mach 3 there would be no available thrust left at all. This installation would also have huge aerodynamic drag, due to air spilling over the intake lip.
To minimise all these losses, a convergent/divergent intake is usually used for supersonic aircraft, but unless this can be made to adjust to varying engine demand and Mach number changes, this intake will be efficient at one Mach number only, and poor flow/efficiency will result at all ‘off design’ Mach numbers. (Lockheed seem to have done an incredible job with the fixed inlets on the F22 Raptor however). Designing a variable inlet in itself is not too difficult, but if you want a design with maximum possible efficiency (no reheat or afterburning) together with totally automated surge protection and operating stability, the task is truly daunting, and before Concorde quite frankly not achieved anywhere.
What any convergent/divergent intake achieves is to use a series of relatively weak oblique shocks to progressively slow the intake air down (Oblique shocks ALWAYS produce supersonic downstream airflow) the unavoidable normal shock is designed to be as weak as possible, and should occur as close as possible to the narrow throat of the intake. The now subsonic air will progressively slow down as it travels through the divergent section of the intake, up to the compressor face. Any time that intake matching is not perfect, large losses quickly occur, with air spilling over the lip of the inlet, and surge/unstart also likely to occur if things go too far off song. (The intake system of Concorde actually sensed the position of the normal shock, and allowed it’s perfect placement by varying the intake surface).
Due to the onset of writer’s cramp/mental fatigue coupled with a desperate need for beer, this will have to do for now guys, I just hope it makes it all a little less ‘clear as mud’
Thanks for your effort. Clearly the answer rests on the interpretation of the axial pressure profile; not as simple to explain as a momentum exchange in a rocket engine or mass flow at subsonic speeds.
Wonder how much of this was learned from the use of the 593s on the 1950s Victor bombers albeit these were not supersonic?
Presume these figures are not comparable, if Olympus reckoned to be the most efficient of the dry supercruise engines - think this discrepancy is that J58 is fundamentally a wet ramjet in cruise, and a smaller core? (OLy usually quoted as the largest core fo any gas turbine?)
The J58 intake and exhaust inlets and outlets create 83% of total trust at M3.2 at 80,000 ft. It first flew in 1963 powering the YF-12A, precursor to the SR-71.
The J58 is a hybrid jet engine: effectively, a turbojet engine inside a fan-assisted ramjet engine. This was required because turbojets are inefficient at high speeds but ramjets cannot operate at low speeds. To resolve this, the airflow path through the engine varied, depending on whether ramjet or turbojet operation was more efficient, thus the term variable cycle. To create this effect, at speeds over 2000 mph the nose cone of the engine was pushed about 2 inches forward to improve the air flow in the ramjet cycle.
Air is initially compressed and heated by the shock wave cones, and then enters 4 stages of compressors, and then the airflow is split: some of the air enters the compressor fans (core-flow air), while the remaining flow bypasses the core to enter the afterburner. The air continuing through the compressor is further compressed before entering the combustor, where it is mixed with fuel and ignited. The flow temperature reaches its maximum in the combustor, just below the temperature where the turbine blades would soften. The air then cools as it passes through the turbine and rejoins the bypass air before entering the afterburner.
At around Mach 3, the initial shock-cone compression greatly heats the air, which means that the turbojet portion of the engine must reduce the fuel/air ratio in the combustion chamber so as not to melt the turbine blades immediately downstream. The turbojet components of the engine thus provide far less thrust, and the Blackbird flies with 80% of its thrust generated by the air that bypassed the majority of the turbomachinery undergoing combustion in the afterburner portion and generating thrust as it expands out through the nozzle and from the compression of the air acting on the rear surfaces of the spikes.
Wonder how much of this was learned from the use of the 593s on the 1950s Victor bombers albeit these were not supersonic?
The Handley Page Victor Mk1 was powered by four Armstrong Siddely Saphire turbojets, and was underpowered like hell. The Mk2 was powered by four Rolls Royce Conway turbofans, was a far more capable aircraft and was the fastest of all the V Bombers. The Avro Vulcan Mk1 was powered by four Bristol Siddely Olympus turbojets, the Mk1 by four Olypus Mk1 1 engines, rated at 11,00lb static thrust and the Mk2 by four 22,00lb thrust Olympus 301 engines. Now allthough the Olympus 593 shares the same name and twin spool turbojet layout, they were in reallity light years apart. The 593, although a development of the earlier engine became an almost total re-design, and was a direct development from the Olymus 320 engine, powering the polititian murdered but absolutely superb BAC TSR2. Machaca Thanks for the superb diagrams for the J58 powerplant. The SR71 (one of my top 3 ever favourite aircraft) was without doubt Kelly Johnson's finest creation, and still remains on the record books as the fastest conventional eaircraft ever built. The axisymmetric intake is in fact a potentially more efficient design than the two-dimentional intake used on Concorde, and is really essential for any Mach 3+ design. (The Mig 25 used a two dimentional intake, but was a crap design that only achieved very brief high speeds by use of brute foce and ignorance). There are two fundimental problems with axisymetrical inlets, that of unstart and also instabily in sideslip conditions. There were many sideslip induced unstart events with the SR71, even aircraft losses occured as a result. There was a club in the SR71 community, known as the 'split helmet club'. This was where as an intake surged or unstarted, as a result of the violent yawing the crew member's bone dome would strike the side of the canopy violently and crack open. To be a member this had to have happened to you. (There were MANY members). I read an article by an SR71 test pilot saying that an unstart was 'like being in a train wreck'. So you can see that this effect was not really desirable for a passenger aircraft carrying 100 passengers). For a note on unstart we have here another extract from the Concorde Air Intake Control System:
Occurs when the intake shock system is expelled from the intake resulting in almost instant engine surging due to enormous flow distortion. Merely throttling the engine will not resolve the problem, as the shock system will not re-establish itself without movement of the intake surfaces. The phenomenon has plagued almost every variable geometry intake ever flown, with the definite exception of this one.
For these reasons an axisymmetric design was ruled out for Concorde, and it was deemed that the inlet had to be both more or less immune from the effects of sideslip, as well as being a self starting design. (Originally the control laws were tweaked to compensate for side slip disturbences, but eventually an aerodynamic solution was found, and side slip signalling/compensation was relegated to blank lines of code in the system software). But in spite of the above, I have nothing but respect and admiration for the SR71.
Bear with me Dude, I think your drawing in post #12 has done it for me, but I'm prepared to be embarrassed yet again.
Its a bypass? The "power" is generated by the action of shock waves slowing the air mass and passes around the 593 to mix with the core engine exhaust within the buckets? So, it is the opposite of a High Bypass fan, in that a core of 'relatively low' velocity air from the 593, is surrounded by a tube of very high velocity/energy air from the intake?
I refer to the 593, buckets and intake as components, because the whole assembly is 'The Engine'. Is that how you see it? Please say I've got it, because its been doing my head in since the statistic (75% thrust from intake) was first mentioned.