Concorde engine intake "Thrust"
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It was so simple I replicated it (as a teenager) in a washbasin....
It is very much like the long-spun-out analogies we tend to use to describe the "sound barrier", and shock waves and the sonic boom, like dropping a stone in a pond, then moving a stick along a water surface at various speeds, then looking at a skiff in a canal, and looking at the waves hitting the bank.
But that simple gadget "showed it all".... even the differences between thin and thick wings by placing different 2D shapes in the stream.
Since then, for me the "sound barrier" was no longer a mystery, even if I didn't learn the maths until much, much later.
Oh, and it also told me, even then, about the difficulities of building transsonic windtunnels....
CJ
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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.
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 ?
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 
Last edited by M2dude; 19th Sep 2010 at 16:44.
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It was a very simple way of demonstrating subsonic and supersonic flow by way of a 2D analogy of wave effects in a very thin sheet of water running down a more-or-less inclined sheet of glass.
Another useful water analogy is the hydraulic jump.
Use the tap/faucet in your kitchen sink to impact a jet of water onto the smooth, horizontal sink surface. The water jet radiates as a thin film of water until it enters the hydraulic jump region. Here, there is a flow discontinuity with the ring of the jump being analogous to a normal shock wave in a compressible gas. Plenty of pictures on the net of jumps.
Main generic things to keep in mind with intake flow -
(a) a normal (perpendicular to the flow direction) shock wave is bad news for flow parameters with lots of losses and so forth
(b) you have a normal shock as you transition through the transonic region whether you like it or not
(c) the trick is to use a series of oblique shocks ahead of the normal shock to step the flow changes progressively so that the abrupt changes across the final normal shock are reasonably minimised. This is the artform inherent in intake design and geometry.
(d) while doing the bits in (c), the designer can pick up lots of useful pressure distributions on the tinwork which is what the intake thrust discussions are all about.
Another useful water analogy is the hydraulic jump.
Use the tap/faucet in your kitchen sink to impact a jet of water onto the smooth, horizontal sink surface. The water jet radiates as a thin film of water until it enters the hydraulic jump region. Here, there is a flow discontinuity with the ring of the jump being analogous to a normal shock wave in a compressible gas. Plenty of pictures on the net of jumps.
Main generic things to keep in mind with intake flow -
(a) a normal (perpendicular to the flow direction) shock wave is bad news for flow parameters with lots of losses and so forth
(b) you have a normal shock as you transition through the transonic region whether you like it or not
(c) the trick is to use a series of oblique shocks ahead of the normal shock to step the flow changes progressively so that the abrupt changes across the final normal shock are reasonably minimised. This is the artform inherent in intake design and geometry.
(d) while doing the bits in (c), the designer can pick up lots of useful pressure distributions on the tinwork which is what the intake thrust discussions are all about.
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Loved your analogy John, and agree 100% with your far more elequent than my 'shockwave summations'. (As you say, you get the normal shock whether you like it or not).
But you sir are in trouble with my wife; I was just demonstrating your water analogy (basically showing off again) and ended up soaking her, and drowwned the kitchen floor.
Dude
But you sir are in trouble with my wife; I was just demonstrating your water analogy (basically showing off again) and ended up soaking her, and drowwned the kitchen floor.
Dude
Last edited by M2dude; 15th Sep 2010 at 07:24.
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ended up soaking her, and drowwned the kitchen floor.
Oh dear ...
.. however, as with the majority of good men of long married character and self preservation knowledge, I'm sure that you grovelled in the approved manner, wiped up the mess ... and the universe returned to normality ...
Oh dear ...
.. however, as with the majority of good men of long married character and self preservation knowledge, I'm sure that you grovelled in the approved manner, wiped up the mess ... and the universe returned to normality ...
I hope that MOST people here find the above descriptions useful and interesting; to me it is one mind-blowing subject.
...and it illustrates that the Great Art of Aerodynmics is Experimental
a little twist here a little waxing and polishing,..., a little area cut away from there...and it---- Still Flies 
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Some Oil Lamps And Diesel Oil
John tullamarine
Oh yes, I found that although only relatively recently married, the art of being humble still prevails. 
Pugilistic Animus
Thanks very for your comments Pugilistic, in my personal opinion the air intake design is one of the most fascinating parts of the whole Concorde tapestry.
If I may, I would now like to mention the 'some oil lamps and diesel oil' story. This is a true story told to me by Dr Ted Talbot, the father of the Concorde Intake, brilliant aerodynamicist and all round amazing gentleman. Ted had been invited in 1975 to speak to the US test pilots at Edwards Air Force Base in California, and after he landed he was invited to take a tour through one of the top secret hangars there, and in this hangar were a few glistening Mach 2.5 design B1A development aircraft. Now Ted had heard that Rockwell were having major difficulties with the engine intakes, and obviously had more than a passing interest in such things, and was allowed to take a close look. Just above and slightly forward of each intake he observed several beautiful made precision total pressure probes mounted under the wings, and although he had a good idea what they were for, said nothing at the time.
That evening, Ted gives his presentation speech to the assembled Test pilots, explaining in fair detail how the Concorde engine intake operated, and that the fact that unlike most other supersonic designs, the engine power was more or less freely variable at Mach 2 and above, even to the extent that if necessary the throttle could be closed all the way to the idle stop. There allegedly many gasps of amazement and disbelief in the room at this, and one B1A pilot was heard to ask his boss 'why the hell can't WE do that John'?. (It should be borne in mind here that the 'traditional' way of slowing down Mach 2+ aircraft is not to touch the throttles initially, and just cut the afterburners. If you don't do it this way many designs will drive into unstart and even flame-out).
After the audience had asked Ted several questions about Concorde, Ted was then invited to ask the assembled USAF and Grumman personnel about the B1A programme, which would be honestly answered within the confines of security considerations. Ted said that he only had one real point to raise; 'I see that you are having major difficulties with wing boundary level interference at the engine inlets'. There was now a gasp of horror from various members of the USAF entourage, 'That's top secret, how the hell do you know that?'. Ted chortled 'it's easy, I saw that you have a multitude of precision pressure sensors under the wing forward of the intakes, that I assume are to measure the wing boundary flows'. Ted then unhelpfully comes up with 'Oh, and you've got the design completely wrong, your intakes are mounted sideways, and that allows the intake shocks to rip into the wing boundary layer, which will completely screw up your inlets at high supersonic speeds. That in my opinion is where most of your problems lie, with wing boundary level interference, but I think that your probes for measuring boundary layer are beautiful, we never had such things'. According to Ted there was not so much uproar at the meeting as much as horror and amazement that this (even then) quite senior in years British aerodynamicist had in a few seconds observed the fundamental design flaw in an otherwise superb but top secret aircraft, and could even see what they were trying to do about it. Ted was asked, 'so you had no boundary layer issues with Concorde then?' Oh we had a few, mainly with the diverter section mounted above the intake' replies Ted, 'but we sorted out the problems relatively easily. 'You said that you did not use precision pressure probes under the wing to measure boundary layer flow fields, so what DID you use then?', asks a Rockwell designer. 'Some oil lamps and diesel oil' replies Ted. The room is now filled with laughter from all those assembled, but Ted shouts 'I am serious, it's an old wind tunnel trick. You mix up diesel oil with lamp black, which you then paint over the wing surface forward of the intakes, where it forms a really thick 'goo', which sticks like glue to the wing'. The pilots in particular seem quite fascinated now, and Ted goes on; 'You fly in as cold air that you can find (we flew out of Tangiers and Casablanca) and flew as fast as you could. As the skin temperature increases with Mach number, the diesel and lamp black 'paint goo' becomes quite fluid, and start to follow the boundary layer flow field. You then decelerated as rapidly as possible, and the flow field 'picture; is frozen into the now again solid 'goo'. After we landed we just took lots of pictures, repeated the process for several flights until we know everything that we needed to know about our difficulties. After doing some redesign work we then repeated the exercise again several times, eventually proving that we'd got things right'. The audience asked Ted if this technique might help them with the B1A, but he replied that although it might help them with accurately illustrating the problem, in his opinion it was irelevant, 'because the intakes are the wrong way round'.
The B1A intake problems were never resolved, and in 1977 the project was cancelled, due to performance and cost issues. However the project was reborn as the B1B, not entering service until 1986. Although an amazing aircraft, with astonishing low altitude performance and capability, it is a fixed intake design, limited to Mach 1.6 at altitude. Ted was right it seems.
Dude
. however, as with the majority of good men of long married character and self preservation knowledge, I'm sure that you grovelled in the approved manner, wiped up the mess ... and the universe returned to normality ...
Pugilistic Animus
Thanks very for your comments Pugilistic, in my personal opinion the air intake design is one of the most fascinating parts of the whole Concorde tapestry.
If I may, I would now like to mention the 'some oil lamps and diesel oil' story. This is a true story told to me by Dr Ted Talbot, the father of the Concorde Intake, brilliant aerodynamicist and all round amazing gentleman. Ted had been invited in 1975 to speak to the US test pilots at Edwards Air Force Base in California, and after he landed he was invited to take a tour through one of the top secret hangars there, and in this hangar were a few glistening Mach 2.5 design B1A development aircraft. Now Ted had heard that Rockwell were having major difficulties with the engine intakes, and obviously had more than a passing interest in such things, and was allowed to take a close look. Just above and slightly forward of each intake he observed several beautiful made precision total pressure probes mounted under the wings, and although he had a good idea what they were for, said nothing at the time.
That evening, Ted gives his presentation speech to the assembled Test pilots, explaining in fair detail how the Concorde engine intake operated, and that the fact that unlike most other supersonic designs, the engine power was more or less freely variable at Mach 2 and above, even to the extent that if necessary the throttle could be closed all the way to the idle stop. There allegedly many gasps of amazement and disbelief in the room at this, and one B1A pilot was heard to ask his boss 'why the hell can't WE do that John'?. (It should be borne in mind here that the 'traditional' way of slowing down Mach 2+ aircraft is not to touch the throttles initially, and just cut the afterburners. If you don't do it this way many designs will drive into unstart and even flame-out).
After the audience had asked Ted several questions about Concorde, Ted was then invited to ask the assembled USAF and Grumman personnel about the B1A programme, which would be honestly answered within the confines of security considerations. Ted said that he only had one real point to raise; 'I see that you are having major difficulties with wing boundary level interference at the engine inlets'. There was now a gasp of horror from various members of the USAF entourage, 'That's top secret, how the hell do you know that?'. Ted chortled 'it's easy, I saw that you have a multitude of precision pressure sensors under the wing forward of the intakes, that I assume are to measure the wing boundary flows'. Ted then unhelpfully comes up with 'Oh, and you've got the design completely wrong, your intakes are mounted sideways, and that allows the intake shocks to rip into the wing boundary layer, which will completely screw up your inlets at high supersonic speeds. That in my opinion is where most of your problems lie, with wing boundary level interference, but I think that your probes for measuring boundary layer are beautiful, we never had such things'. According to Ted there was not so much uproar at the meeting as much as horror and amazement that this (even then) quite senior in years British aerodynamicist had in a few seconds observed the fundamental design flaw in an otherwise superb but top secret aircraft, and could even see what they were trying to do about it. Ted was asked, 'so you had no boundary layer issues with Concorde then?' Oh we had a few, mainly with the diverter section mounted above the intake' replies Ted, 'but we sorted out the problems relatively easily. 'You said that you did not use precision pressure probes under the wing to measure boundary layer flow fields, so what DID you use then?', asks a Rockwell designer. 'Some oil lamps and diesel oil' replies Ted. The room is now filled with laughter from all those assembled, but Ted shouts 'I am serious, it's an old wind tunnel trick. You mix up diesel oil with lamp black, which you then paint over the wing surface forward of the intakes, where it forms a really thick 'goo', which sticks like glue to the wing'. The pilots in particular seem quite fascinated now, and Ted goes on; 'You fly in as cold air that you can find (we flew out of Tangiers and Casablanca) and flew as fast as you could. As the skin temperature increases with Mach number, the diesel and lamp black 'paint goo' becomes quite fluid, and start to follow the boundary layer flow field. You then decelerated as rapidly as possible, and the flow field 'picture; is frozen into the now again solid 'goo'. After we landed we just took lots of pictures, repeated the process for several flights until we know everything that we needed to know about our difficulties. After doing some redesign work we then repeated the exercise again several times, eventually proving that we'd got things right'. The audience asked Ted if this technique might help them with the B1A, but he replied that although it might help them with accurately illustrating the problem, in his opinion it was irelevant, 'because the intakes are the wrong way round'.
The B1A intake problems were never resolved, and in 1977 the project was cancelled, due to performance and cost issues. However the project was reborn as the B1B, not entering service until 1986. Although an amazing aircraft, with astonishing low altitude performance and capability, it is a fixed intake design, limited to Mach 1.6 at altitude. Ted was right it seems.
Dude
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Hi Ventus45
There was no real 'magic' as far as the intake mounting to the wing goes. There were just four heavy duty attachments as well as a sliding transition ring between the intake and engine compressor case. The attachment links on the intake itself were also allowed to move fore and aft a little The two rearmost attachment points ran along the centre line of the intake assembly and were totally enclosed within a large arrow shaped diverter, the large seal of which butted up against the wing lower surface.
Dude
There was no real 'magic' as far as the intake mounting to the wing goes. There were just four heavy duty attachments as well as a sliding transition ring between the intake and engine compressor case. The attachment links on the intake itself were also allowed to move fore and aft a little The two rearmost attachment points ran along the centre line of the intake assembly and were totally enclosed within a large arrow shaped diverter, the large seal of which butted up against the wing lower surface.
Dude
I don't even have further comments that story is unbeatable---boundary layer mapping with lampblack and diesel fuel vs high tech probes...there must be a million How many aerodynamiacists does it take to----jokes for that one
M2dude
Thanks for relaying that anecdote---LOL
M2dude
Thanks for relaying that anecdote---LOL
Guest
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The SR-71 was developed in 1959-60. Cocktail napkins and slide rules. Johnson drank Scotch and water, the Concorde is more than beautiful, she is a totem of genius.
She also blazed trails the Blackbird never had to face, but credit where credit is due.
Mr. Pratt, Mr. Whitney, Ampersand.
M2dude that story is priceless, I envy your experience with CONCORDE
bear
She also blazed trails the Blackbird never had to face, but credit where credit is due.
Mr. Pratt, Mr. Whitney, Ampersand.
M2dude that story is priceless, I envy your experience with CONCORDE
bear
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bearfoil
Absolutely credit where it is due. The YF12/SR71 was without doubt Kelly Johnson's finest creation. (I'll let you into a secret, it was the SR71 inlet that first got me 'hooked' into the world of shockwave management).
As I said in a previous post, you will find nothing but respect in the family of Concorde for the SR71.
Dude
She also blazed trails the Blackbird never had to face, but credit where credit is due.
As I said in a previous post, you will find nothing but respect in the family of Concorde for the SR71.
Dude
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At the 1974 Farnborough airshow, the crew of the SR-71 that had just done a record New York-to-London flight, was treated to a flight in Concorde.
Since they could hardly "pay back in kind", they reciprocated by treating us all at Fairford to our own "private airshow".
I've never forgotten that.... seeing that alien shape in the sky, that already had been flying routinely at Mach 3+ before Concorde first flew.
CJ
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I remember it well CJ; I seem to remember that the guys flew in pre-production aircraft 101 (G-AXDN), and were full of praise about Concorde and the fact that they flew for the first time at 60,000' WITHOUT a pressure suit. (And in the case of SR71 crews this was more akin to a full blown spacesuit).
Dude
Dude
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OMG I regularly went up to 63,000' on test flights, without too much blood bubbling I'm pleased to report. The physiological effects of altitude was well known to most of us, and dictated just about everything about the design philosphy for the environmental systems on Concorde.
These test flight excursions at 63,000 were when we used to check the intake surge margins by doing a near zero G pushover at Mach 2 from a zoom climb. (Which I mentioned in another Concorde thread). At the top of the 'bunt' the throttles would be pulled back, to prevent overspeeding; at the point of zero alpha the local Mach number at the intake face would be far higher than any other time, as well as wing flow distortion would also be at maximum, so hence the surge margin check. With all this going on you really did ask yourself if this was really an airliner.
Dude
These test flight excursions at 63,000 were when we used to check the intake surge margins by doing a near zero G pushover at Mach 2 from a zoom climb. (Which I mentioned in another Concorde thread). At the top of the 'bunt' the throttles would be pulled back, to prevent overspeeding; at the point of zero alpha the local Mach number at the intake face would be far higher than any other time, as well as wing flow distortion would also be at maximum, so hence the surge margin check. With all this going on you really did ask yourself if this was really an airliner.
Dude
Last edited by M2dude; 18th Sep 2010 at 15:31.
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This thread reminds me of when I was a young airman (Airframes) & I asked some engine guys "Where does the thrust act in a jet engine". They can't have been very bright because they just told me the 'Change of momentum' business.
I then got hold of a flying AP & it had a diagram of the forward & rearward gas loads in a RR Avon engine (With & without reheat) that cleared things up for me.
It would be interesting to see all the forward & rearward gas loads in the Concorde intakes, engines, nozzles ect, after all, something is 'Pushing' somewhere, it just needs to be split down into sections to see where.
I can understand manufacturers having commercial secrets, maybe that's why they don't publish this information on the internet, but it all boils down to simple forward & rearward gas loads somehere in the works.
I then got hold of a flying AP & it had a diagram of the forward & rearward gas loads in a RR Avon engine (With & without reheat) that cleared things up for me.
It would be interesting to see all the forward & rearward gas loads in the Concorde intakes, engines, nozzles ect, after all, something is 'Pushing' somewhere, it just needs to be split down into sections to see where.
I can understand manufacturers having commercial secrets, maybe that's why they don't publish this information on the internet, but it all boils down to simple forward & rearward gas loads somehere in the works.
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I have often wondered about depressurisation in Concorde.
Above 40,000ft involves pressure breathing if no pressure suit.
Was there a provision for flight deck crew to pressure breath in event of explosive decompression?
It was speculated in my av med course that Concorde was designed in a way that "explosive" decompression was not taken into account but we had no real idea or information.
John
Above 40,000ft involves pressure breathing if no pressure suit.
Was there a provision for flight deck crew to pressure breath in event of explosive decompression?
It was speculated in my av med course that Concorde was designed in a way that "explosive" decompression was not taken into account but we had no real idea or information.
John