ChritiaanJ
My point was we had not shown just how complex and difficult the Concorde intake aerodynamics were in these posts. I have mentioned NOTHING about the complexities of the generation of the generation of the shock system as I thought it might be a little 'heavy' in the context of this topic, but in defference to you, maybe I will for the benefit of everyone
ELSE here:
The above diagram shows a broad view of the intake at Mach 2 cruie. What is not shown here is, if you like, the 'very first' shock; this comes off the wing leading edge, reducing the local Mach number (Mo) to around Mach 1.9 for an inner intake, Mo for the outer intakes is a little lower.
Here are some extracts from The Concorde Air Intake Control System.. You may also want to refer to my previous 'whole powerplant' diagram:
[quote]
Assume Mach 2 supersonic cruise conditions, with the intake operating critically. Underwing local Mach Number is assumed to be Mach 1.9 (a good average for the inboard and outboard intakes). The ramp angle is assumed to be 9.5 degrees (about 45% on the Manual Control Panel's ramp position indicator). As the entry airflow enters the intake it encounters the 1st shock, which at normal Mach Numbers is just forward of the cowl lip. As well as the air experiencing a reduction in velocity, it is turned downwards to follow the profile of the fixed (7 deg') wedge compression surface. The Mach Number at this point has now fallen to approx' Mach 1.65. As the shock is not 'on lip' there is a small amount of airflow lost over the lip known as 'Supersonic Forespill', this generating moderate losses in the way of form drag etc. In fact the losses incurred by this spill drag equates to about a tonne of fuel burnt (or a corresponding reduction in payload), but to allow the intake to cope with aircraft Mach overshoots, without surging this unfortunately is a necessary evil.
As the airflow meets the 2nd and 'Fan' shocks, it is subjected to further turning down, following the forward ramp profile, which produces a 5.75 deg' total turn-down by the bottom of the ramp. (So the air is subjected to the initial 7 degree turn down plus a turn down that depends on the actual ramp angle and a 5.75 deg' turn down imposed by the curve in the ramp profile). The Mach Number after the second shock has fallen to approx' Mach 1.57, and after the final stage of the fan shock to approx' Mach 1.37. Transition of the airflow through the fan shock produces a staccato increase in Ps and reduction in velocity. What is particularly interesting about this process, known as 'isentropic turning', is that there is absolutely NO LOSS in Pt (Total pressure) as a result, making the utilisation of an isentropic fan shock an extremely efficient way of carrying out the compression process. As the downward inclined airflow meets the cowl lip, which itself is inclined upwards at 12 deg's, the 4th shock is formed. Because of the relatively low local Mach Number at this point (M1.37) and the fairly shallow approach angle of the airflow relative to the cowl lip (3.25 deg's, see below), a strong oblique shock is produced. This shock is inclined upwards towards the bleed slot (the gap between the ramps) and this slot has the effect of modifying the shape of this shock into a gentle curve, the upper component of this shock helps force the secondary airflow into the bleed gap. The total airflow turndown at this point now is the initial 'fixed wedge' 7 deg's plus the combined turndown as a result of the 9.5 degree ramp angle, and the 'isentropic turn' of 5.75 degrees]. We therefore at this point experience a total turndown of 7 + (9.5 - 7) + 5.75 = 15.25 degrees]. (As the ramp angle is taken relative to the local horizontal and not the 7 degree wedge, we subtract 'wedge angle' from ramp angle). This airflow then, at an incident angle of 15.25 degrees relative to the horizontal. The approach angle of the airflow onto the cowl lip is therefore 15.25 - 12 = 3.25. (This producing our nice 'strong oblique' shock rather than a normal shock). Our oblique shock has the effect of starting to turn the airflow back into line with the engine, in fact to within about 5 degrees] of the local horizontal.
Now for some real confusion; Although we have produced an oblique shock, as far as the local airflow at the base of this shock is concerned, a small amount of the shock is in fact normal and we therefore end up with a mix of just supersonic air (upper region) and just subsonic air (lower region). In fact, because of the curved nature of the shock, we end up with a progressively varying mix of Mach Numbers in the downstream section. As a result of the coalescing of these supersonic/subsonic airflows, we end up with a few very weak near normal shocks that radiate rearwards from the 4th shock, these shocks collectively being known as ‘the terminal shock’. The terminal shock is about half intake height and stands over the bleed slot and can be considered as a ‘virtual’ single weak normal shock. The downstream airflow is now mixed and finally subsonic, having fallen to about Mach 0.98. ]Beyond the terminal shock, the subsonic (only just) airflow continues its journey to the engine, through the divergent (diffuser) section of the intake. As well as functioning as a conventional subsonic diffuser (as the airflow passes through the duct, it's velocity progressively reduces and it's static pressure simultaneously increases), this section also has the effect of causing the primary flow to turn the final 5[/font][FONT='Arial','sans-serif']o[/font][FONT='Arial','sans-serif'] back into line with the engine. As far as the primary airflow is concerned it has now come to the end of its journey to the engine face, but before we deal with the secondary airflow, we now have to dispel a little Concorde folklore:
Contrary to popular belief, MN1 engine compressor face Mach Number) has NOTHING directly to do with intake operation as such, being entirely dependent on engine mass flow and compressor face cross sectional area. If the intake goes 'off tune' for any reason, MN1 remains the same, only the losses incurred in the course of producing that Mach Number would increase markedly. Even if the intake ‘wasn’t there’ this Mach Number would still be the same. (There would be a massive normal shock across the face of the compressor and probably barely enough P1 left to produce any real thrust at all). As far as our intake is concerned, at the compressor face and assuming 'design' engine mass flow, the engine airflow MN1, will be at Mach 0.49.
The now subsonic secondary airflow passing over the rear ramp is channelled to the four secondary air doors by some carefully designed cascade ducting. The secondary flow now finally completes its journey by travelling through the engine bay as cooling air and exiting via the gap between the primary nozzle and the secondary nozzle structure. This air is now used to give the rapidly expanding exhaust flow a relatively high pressure cushion and so limit this expansion, reducing 'flaring' of the exhaust efflux and hence the massive potential loss in thrust. Together with the divergent nozzle of the open secondary nozzle buckets, the secondary airflow helps to maximise nozzle thrust
Local sensing of under-wing airflows is not practical, in termes of accuracy and predictability, and so local manometric data was used to accurately synthesise the flow field conditions, and the use of only one internal intake static pressure tapping was required to accurately predict the precise shock system geometry.
So we can see tha there is nothing at all simple about creating this amazing shockwave cocktail, and the control of all this was also something else, and if we go off song even slightly, then reduction in powerplant efficiency and/or surge will result.
In the ideal world, our intake would just operate in a critical manner, but superimposed on this 'performance requirement' are limitations placed on the control pressure ratio, the variable limits for maximum and minimum ramp angles, as well as maximum engine mass-flow demand. All of these variables change with intake local Mach number; the intake acually limiting engine N1 at high Mach number, low temperature conditions. Oh, and changes of aircraft incidence have also to be instantly compensated for, particularly at very low Alpha. Incidence will both alter the capture airflow AND affect intake local Mach number.
I hope that MOST people here find the above descriptions useful and interesting; to me it is one mind-blowing subject.
The 'oil lamps and diesel oil' story in a future post, and no ChristiaanJ, it's not just about Casablanca, perhaps you will allow me to explain ?
Dude
