Lowering stall speed

[Dafydd Llewellyn]

Since most aeroplanes are elevator-limited at their most forward centre of gravity, modifications that are aimed at increasing the wing lift are generally ineffective in reducing the stall speed. Further, anything that increases the nose-down pitching moment of the wing will increase the download at the tail, so if the aircraft is - as is normally the case - elevator limited, increasing the span of the flaps (or anything else that increases their power) will generally act to INCREASE the stall speed. Jabiru tried this way back in their early days, with exactly that result. Message: The stall speed is a consequence of the design of the whole aeroplane, not just the wing or the flaps.

 

 

[metalman]

Dafydd, you had a fair bit to do with the Skyfoxs development, did they try the drooping flaperons ( like the kitfoxs) in the early stages? If so did it lower the stall much ? And how far did they droop them ( if they did infact try it) , I've flown a few planes with drooping ailerons and the loss in control deflection is more annoying than the higher stall speed,in my opinion.

 

Matty

 

 

[Head in the clouds]

This has now become a classic demonstration of how some people allow 'Perfectionism to impede Completion' as so often becomes the case when some people set out to design and build their own plane. In the real world, however, a lot of things have been demonstrated to work acceptably well where the theoreticians have categorically declared that they can't possibly do so.

 

I don't think anyone suggested that flapperons were the perfect scenario but the fact is they have been demonstrated to work acceptably well for people with reasonable flying skills. Quite possibly they may not be a great idea for some tricycle drivers with dead feet.

 

One of their great advantages is that they can be retro-fitted relatively easily to aircraft with full span ailerons with some benefit and little or no other modification to the airframe. Those same aircraft can benefit from then being able to have an extra setting reflexing the ailerons and often gain a useful increase in cruise speed where the structure allows it, the Drifter being a good example of that, or if not a speed increase then a fuel saving/efficiency gain.

 

But - the point is, the OP wasn't trying to get a 40% decrease in stall speed*, he was actually chasing just a knot or two and around 5 degrees of aileron droop would achieve that easily enough with little or no noticeable change in handling characteristics, the KR2 is horrible enough as it is, it'd take a lot to make it worse, but that's just my opinion of course.

 

*Where can I find a reference to that Bob? I have all the NACA Langley Reports, Notes, Memoranda, WRs etc 1917-'58 and have done a quick search and haven't come across anything in them that suggests achieving that with multiple slotted, let alone single slotted.

 

 

[metalman]

If he wants to "make" the numbers look right , bump the idle up a couple of hundred rpm's, that will hold the stall off a tad

 

 

[rgmwa]

Thanks Bob. I'll have to read that a few times to get my head around it. Interestingly the RV10 has reflexive flaps for reduced cruise drag and a few more knots. The RV-12's flaperons have (I've read, but not sure how accurate this is) about +24/-12 deg movement at zero droop and +18/-11 at the max droop of 30 deg. Van's lead designer was Ken Kreuger.

 

rgmwa

 

 

[Dafydd Llewellyn]

Dafydd, you had a fair bit to do with the Skyfoxs development, did they try the drooping flaperons ( like the kitfoxs) in the early stages? If so did it lower the stall much ? And how far did they droop them ( if they did infact try it) , I've flown a few planes with drooping ailerons and the loss in control deflection is more annoying than the higher stall speed,in my opinion.Matty

Yes, they did try the drooping flailerons - the initial configuration of the Skyfox when I first flew it, was a Kitfox 3 with an aeropower engine; it had, like the K3, the drooping ailerons, operated by what is now the trim lever. They were completely ineffective in reducing the stall speed, because the aircraft was elevator-limited - i.e. at forward CG, one hit the up-elevator stop before the wing stalled; so increasing the lifting capability of the wings was actually counter-productive, because in doing so, the droop increased the nose-down pitching moment of the wing, which further exacerbated the inadequacy of the tail. That could not be fixed by increased elevator travel, because the "flat-plate" tail surfaces were already at the limits of their capability, due to the limited tailplane span necessary to allow the aircraft to be road-towed with its wings folded. We made the greatest improvement (one knot) in stall speed by gap-sealing the elevators, and that was just sufficient to meet the 40 KCAS requirement of CAO 101.55.

Further, putting the "flaps" down loaded-up the aileron circuit so that it required considerable effort to turn onto final - and the movement of the flailerons in so doing caused the droop to return to zero. The mechanism was obviously in need of complete re-design if that feature was to function properly. So the decision was made to delete the droop function and use the lever for the elevator trim instead.

 

 

[Dafydd Llewellyn]

If he wants to "make" the numbers look right , bump the idle up a couple of hundred rpm's, that will hold the stall off a tad

That effect is a clear indication that the aircraft is elevator-limited; the effect of a couple of hundred revs is mainly to increase the velocity over the tailplane (unless it's a T-tail, of course). The downside is increased landing float. We also used that trick on the Skyfox, to get it through. All these fine-tuning tricks can gain is quite limited; about 2 ~ 3% at best. Maybe that would suffice for the KR-2 in question.

In a certification scenario, one has to take into account a whole lot of aspects, and find the best trade-off for all of them. The "stickiness" of the Skyfox ailerons when drooped, would have prevented the aircraft from passing the lateral stability requirement - which is that it must tend to raise the low wing when the ailerons are released from a full crossed-control situation, at any available power and speed and centre of gravity (including assymetric fuel).

 

Aircraft design is a complicated juggling act, and the textbooks only discuss one aspect at a time. This is where the "theoreticians" - at least, the amateur ones - tend to come unstuck. In the real World, one has to take account of all of them. Going to extremes in any one area will invariably produce a problem somewhere else.

 

 

[facthunter]

Trying to raise the nose by applying a down force is contra indicated. Short tail moment arms make it worse. Reducing the allowable forward limit of Cof G will help, as will having more wingspan. Fowler flaps work a treat as Cessna's high wing aircraft show. They cost a bit more to make and maintain with tracks etc but there you are. Nev

 

 

[Dafydd Llewellyn]

Trying to raise the nose by applying a down force is contra indicated. Short tail moment arms make it worse. Reducing the allowable forward limit of Cof G will help, as will having more wingspan. Fowler flaps work a treat as Cessna's high wing aircraft show. They cost a bit more to make and maintain with tracks etc but there you are. Nev

Yes, it's a bit discouraging to have to force the tail down - thus increasing the load the wing has to carry - in order to increase the wing lift. This thought leads one to a canard layout. However, that has its own set of disadvantages; overall, the conventional layout works best, unless some special consideration prevails. The harder you try to get a high maximum lift coefficient for the aircraft as a whole, the more the secondary factors make it difficult. The Westland Lysander was quite a good example of that, and so is the Fiesler Storch.

This is why, if one wants to design an aircraft, one needs to understand both the longitudinal stability equation and the longitudinal balance equation - and also know how to take account of the downwash field of the lifting surfaces. Any fool can come up with something that flies - but to get something that flies really well, requires some depth of knowledge. To get a really good result, you have to get into the "flute music". That puts it out of the scope of this website, I believe. A good starting point is Stinton, "The design of The Aeroplane".

 

 

[facthunter]

The Canard appears to be a solution but again it seems to have it's own set of problems. It's invariably a pusher and is often too sensitive in pitch as well as a swag of other bad (and annoying) features. You tend to get back to a fairly conventional plane done right, when you are below about 220 knots top speed. Distribution of the mass along the pitch axis has to be looked at too. If it is well to the extremes it will love to spin (flat). Aircraft with the engine on top of the tail fin are my pet hate.. Have to be an absolute death trap..Nev

 

 

[Dafydd Llewellyn]

The Canard appears to be a solution but again it seems to have it's own set of problems. It's invariably a pusher and is often too sensitive in pitch as well as a swag of other bad (and annoying) features. You tend to get back to a fairly conventional plane done right, when you are below about 220 knots top speed. Distribution of the mass along the pitch axis has to be looked at too. If it is well to the extremes it will love to spin (flat). Aircraft with the engine on top of the tail fin are my pet hate.. Have to be an absolute death trap..Nev

I studied a canard layout as my graduation design project. Even built a wind-tunnel model and tested it. The principal problem with a canard is that longitudinal stability of an aircraft having two lifting surfaces one behind the other, requires that the front lifting surface must always operate at a higher lift coefficient than the rear one. This applies whether the layout is conventional (larger surface in front) or Canard (larger surface to the rear) or tandem-wing (equal size surfaces). The result, for a canard, is that the foreplane has to work harder than the wing; you can't generally put high-lift devices on the wing, they have to go onto the foreplane. So for a given combined area of the two lifting surfaces, the canard uses them less effectively than does a conventional layout. So in general, the maximum lift coefficient for a canard has to be lower than for an equivalent conventional layout - i.e. for the same weight and stall speed, the canard needs a larger total area.

The second big problem with a canard is that the larger lifting surface - which provides the volume for fuel - is not close to the centre of gravity; the CG usually needs to be well forward of the wing. So you cannot put the fuel in the wings, if they are unswept. The Vari-Eze's swept wing is there so the front corners of it can be close enough to the CG to provide fuel stowage.

 

The more one studies the Canard layout, the more the three-surface layout makes sense - especially if the front surface is a free-floating pitch-trim device, but does not contribute to longitudinal stability - e.g. the Piaggio Avanti. Canards also make sense as pitch-trim devices for supersonic flight. However, they are getting to be very complex devices - and the best place for a propeller is NOT behind something that can shed ice.

 

All this is interesting - but rather off-topic.

 

 

[Bob Llewellyn]

This has now become a classic demonstration of how some people allow 'Perfectionism to impede Completion' as so often becomes the case when some people set out to design and build their own plane. In the real world, however, a lot of things have been demonstrated to work acceptably well where the theoreticians have categorically declared that they can't possibly do so.I don't think anyone suggested that flapperons were the perfect scenario but the fact is they have been demonstrated to work acceptably well for people with reasonable flying skills. Quite possibly they may not be a great idea for some tricycle drivers with dead feet.

 

One of their great advantages is that they can be retro-fitted relatively easily to aircraft with full span ailerons with some benefit and little or no other modification to the airframe. Those same aircraft can benefit from then being able to have an extra setting reflexing the ailerons and often gain a useful increase in cruise speed where the structure allows it, the Drifter being a good example of that, or if not a speed increase then a fuel saving/efficiency gain.

 

But - the point is, the OP wasn't trying to get a 40% decrease in stall speed*, he was actually chasing just a knot or two and around 5 degrees of aileron droop would achieve that easily enough with little or no noticeable change in handling characteristics, the KR2 is horrible enough as it is, it'd take a lot to make it worse, but that's just my opinion of course.

 

*Where can I find a reference to that Bob? I have all the NACA Langley Reports, Notes, Memoranda, WRs etc 1917-'58 and have done a quick search and haven't come across anything in them that suggests achieving that with multiple slotted, let alone single slotted.

NACA Technical Note 808 shows the two-dimensional lift coefficient of 23012 with a single 0.3c Fowler flap at 40 degrees deflection, as 3.3 compared to the naked section at 1.52; I think the same data is plotted out in Abbot & Von Doenhoff. This equates to a dCl of ~1.78; at KR-2 reynolds numbers and planform, the maximum practicable root Clmax (unflapped) will be about 1.4 If one assumes ~65% span flaps, with a bit of effective taper, then the fowler in question has the potential to just about double the Clmax (the inner 65% span carries about 80% of the load, giving an overall dCl of 1.78 x 0.8 = 1.42). This would give 1/1.42, or a ~30% reduction in Vs.

NACA Report 723 shows 230XX in three thicknesses, with a moderately undeveloped 40% chord double-slooted flap pack. There is more information therein than meets the eye, as 23012 is a leading-edge staller, whereas 23030 is definitely a trailing-edge staller (23021 changes with Re, but may be considered TE staller for small a/c purposes). In that Report, 23012 had a max dCl of 1.9+; 23030 had ~2.67. As it's the upper rear wing section curvature that is most affected by the flap local suction (thus leading to greater circulation etc), the 23030 data should read across to a thinner aft-cambered TE stalling section, like USAF 35B. On the same basis as the Fowler considered above, the stall speed comes down to 59% of the flapless value - viola!.

 

A conservative aerodynamicist would regard this pasting of the 23030 dCl onto any old rear-loaded airfoil as an optomistic guess, as much because of the retrospectively known vagueness in the NACA correlations between excessive tunnel turbulence and effective reynolds number, as because of the paucity of data on short-bubble behaviour of 23030 (which should greatly affect the translatability of the data to 6-digit sections). It is of note that the 6-digit data published in Abbot & von Doenhoff predates the discovery - by Hoerner - of the VD tunnel's vast excess of turbulence; so the 6-digit Clmaxes given in A&vD are optomistic at least.

 

So - can we get a 40% reduction in Vs, or not? Well, NACA WR L310 - which is subsequent to the 40% double-slotted tests - shows that minor changes in the fore flap - aft flap relationship can give a dCl variation - on a 0.3c flap pack - of 0.56. As this research was based directly on the 0.4c double slotted flap pack, it is reasonable to assume that the 0.4c pack could be improved by as much as 0.3 (dCl) by similar tweaking; which is larger than the various tunnel errors discussed previously, and suggests that the previous generalisation about wide applicability of the 23030 data is defensible.

 

But wait, there's more! An ace lurks up the sleeve, in the form of some data - reported I think in Hoerner's "Fluid Dynamic Lift" - showing a remarkable benefit from a downturned flap TE on a slotted flap. Or perhaps not remarkable - if a plain flappette gives even a 50% improvement on flap nose suction (i.e. local dCl of ~0.6), then the knock-on effect should give a very similar magnitude of benefit to the whole wing / flap pack assembly - i.e. a 50% improvement in overall wing dCl.

 

But is the idea of an overall Clmax of >4 the ramblings of a madman? Schrenk's tests on blown fat airfoils got to a Clmax of well over 5 (from memory). If we regard the rear flappette as increasing the nose suction of the rear flap, which in turn increases the nose suction of the front flap, creating a boundary layer control equivalent to a wing with a suction slot at the TE, it becomes apparent that the overall wing circulation behaviour of a highly flapped wing should be very comparable to that of a 'forced' boundary layer wing.

 

It is of note that the airliners of today, which might be expected to be designed to later data than the published NACA work, are in fact using flap packs that are linear developments of the NACA work, and not very much different at all.

 

 

[Dafydd Llewellyn]

Yes, if you look only at the potential gain in the maximum lift coefficient, there appear to be some startling possibilities. There are, however, several major disadvantages in these extreme possibilities:

 

Firstly, if you look at the effect of such a powerful flap system on the airfoil pitching moment, Abbott & Von Doenhoff show, for example, that the common 26% chord 2h slotted flap at 30 degrees deflection, increases the nose-down pitching moment of 23012 from around -0.02 to around -0.36 - i.e. eighteen times the zero-flap value, for a maximum lift increase of about 90%. So the download on the tail increases much more than the lift on the wing. This effect reduces the potential gain in lift, because the wing must carry the additional download. In practice, this effect reduces the potential benefit of the flaps to around half what the lift coefficient gain suggests it possible. Further, the increased tailplane size necessary adds to the cruise drag. You also need a larger vertical tail, to maintain directional stability - and increased dihedral to offset the spiral instability due to the larger vertical tail.

 

Secondly, the induced drag increases in proportion to the square of the lift coefficient - and there's an addition to the simple induced drag due to the variation of the drag from the flaps; this results in the minimum drag speed being increased; so the tendency towards speed instability ("on the back side of the drag curve") is markedly increased; extreme STOL aircraft are noted for tending to fall out of the air due to this; the Helio Courier etc were notorious for it.

 

Thirdly, the lateral stability is adversely affected because of the effect of the asymmetric slipstream interacting with a powerful flap system, in the crossed-control case required by the certification testing.

 

Fourthly, the behaviour of such aircraft in an aborted landing can be rather startling - in the case of the Helio Stallion, approaching lethal. Generally, the more powerful the flap system, the more powerful needs to be the engine - and the combination often produces a violent pitch-up tendency when power is applied to go around; there are limits set in FAR 23 to the stick force necessary to control this - they are there as a result of the Helio Stallion, which needed both hands on the wheel, so you could not re-trim the thing, because it lacked an electric trim capability. It's all "designing into a corner".

 

All these things can be catered for in the original design - the Twin Otter is a prime example - but there's a price for it. STOL was fashionable in the 1960s, with aircraft like the Helio Courier, turbo Beaver, Pilatus turbo-porter, Caribou, etc, which can show startling performance; however they did not make much market penetration except where special circumstances gave them an advantage. Ansett, for example, found to his cost that a Caribou, when operated under normal civil rules for balanced field length, could not lift as much freight as a DC-3, on identical engines & propellers. (It became known as the Caribou boo). In military use, which ignores the risk of engine failure on takeoff, the Caribou was remarkable. It either works, or you're dead.

 

Cessna found it advisable to remove the fourth flap notch from the 182, because the average pilot was barely able to contend with the side effects of the increased performance. The latest C172 has reduced flap deflection, so it's difficult to make a steep power-off approach in it. I can only assume this is a result of the cost of product liability insurance, but this is the way things are going.

 

 

[facthunter]

Agree with that Daffydd. though the Caribou did have P&W R-2000's against the P&W 1830's in the DC3. The DC3 operated on a PK chart , in PNG which was "developmental" allowing an overweight figure to be used as a concession on take-off. The Ansett DC-3's on the mainland had been converted to Wright 1820 single row engines, and I presume the PNG ones were similar, which were credited with the same power as the P&W, when they may have been a little less. Your point that lift devices require more engine power and involve trim changes is a good point to consider, as well as more complex controllability effects. Nev

 

 

[Bob Llewellyn]

Yes, if you look only at the potential gain in the maximum lift coefficient, there appear to be some startling possibilities. There are, however, several major disadvantages in these extreme possibilities:Firstly, if you look at the effect of such a powerful flap system on the airfoil pitching moment, Abbott & Von Doenhoff show, for example, that the common 26% chord 2h slotted flap at 30 degrees deflection, increases the nose-down pitching moment of 23012 from around -0.02 to around -0.36 - i.e. eighteen times the zero-flap value, for a maximum lift increase of about 90%. So the download on the tail increases much more than the lift on the wing. This effect reduces the potential gain in lift, because the wing must carry the additional download. In practice, this effect reduces the potential benefit of the flaps to around half what the lift coefficient gain suggests it possible. Further, the increased tailplane size necessary adds to the cruise drag. You also need a larger vertical tail, to maintain directional stability - and increased dihedral to offset the spiral instability due to the larger vertical tail.

 

Secondly, the induced drag increases in proportion to the square of the lift coefficient - and there's an addition to the simple induced drag due to the variation of the drag from the flaps; this results in the minimum drag speed being increased; so the tendency towards speed instability ("on the back side of the drag curve") is markedly increased; extreme STOL aircraft are noted for tending to fall out of the air due to this; the Helio Courier etc were notorious for it.

 

Thirdly, the lateral stability is adversely affected because of the effect of the asymmetric slipstream interacting with a powerful flap system, in the crossed-control case required by the certification testing.

 

Fourthly, the behaviour of such aircraft in an aborted landing can be rather startling - in the case of the Helio Stallion, approaching lethal. Generally, the more powerful the flap system, the more powerful needs to be the engine - and the combination often produces a violent pitch-up tendency when power is applied to go around; there are limits set in FAR 23 to the stick force necessary to control this - they are there as a result of the Helio Stallion, which needed both hands on the wheel, so you could not re-trim the thing, because it lacked an electric trim capability. It's all "designing into a corner".

 

All these things can be catered for in the original design - the Twin Otter is a prime example - but there's a price for it. STOL was fashionable in the 1960s, with aircraft like the Helio Courier, turbo Beaver, Pilatus turbo-porter, Caribou, etc, which can show startling performance; however they did not make much market penetration except where special circumstances gave them an advantage. Ansett, for example, found to his cost that a Caribou, when operated under normal civil rules for balanced field length, could not lift as much freight as a DC-3, on identical engines & propellers. (It became known as the Caribou boo). In military use, which ignores the risk of engine failure on takeoff, the Caribou was remarkable. It either works, or you're dead.

 

Cessna found it advisable to remove the fourth flap notch from the 182, because the average pilot was barely able to contend with the side effects of the increased performance. The latest C172 has reduced flap deflection, so it's difficult to make a steep power-off approach in it. I can only assume this is a result of the cost of product liability insurance, but this is the way things are going.

pish tish, when ya pull the flap lever, you just have it move the whole wing fowards a foot or so! (304.8mm for those in the current century); offsets the pitching moment, increases the stability. It may be a structural challenge...

The NACA modification of the F2-F (Buffalo; my memory was in error ie not the F4-F tests), fitted with a full-span 2H single-slotted flaps with both slot-lip and in-flap plain ailerons, gave acceptable handling at an overall dCl up to 0.88, at low speed / flight idle; higher dCls were accompanied by pitch instability and a tendency to flick roll as a stall warning, though the 1943 military test pilots had no trouble handling it...

 

In essence, the F2 tests showed a need for powered controls, artificial force feedback, and fly by wire. But with these minor issues addressed, there is no reason that a 1,000 hp single seat recreational aeroplane to a military design standard could not routinely achieve overall Clmaxes in excess of, say, 3.0.

 

As the KR2 has an excess seat / load capacity, there is no reason it could not be converted to fly-by-wire, using off-the-shelf servos and an industrial controller implementation of a PC for stability/feedback issues. And as the 912/914 is bulletproof (they say), fit a 914 with a microsquirt programmable fuel injector, and nitrous oxide injection into the eye of the turbocharger compressor wheel (about 5lb/minute should give about 220HP WOT). Extend the main legs a bit so the 84" propellor needed clears the ground (it IS a taildragger, of course?), and Bob (no relation) zyouruncle...

 

 

[Bob Llewellyn]

Yes, if you look only at the potential gain in the maximum lift coefficient, there appear to be some startling possibilities. There are, however, several major disadvantages in these extreme possibilities:Firstly, if you look at the effect of such a powerful flap system on the airfoil pitching moment, Abbott & Von Doenhoff show, for example, that the common 26% chord 2h slotted flap at 30 degrees deflection, increases the nose-down pitching moment of 23012 from around -0.02 to around -0.36 - i.e. eighteen times the zero-flap value, for a maximum lift increase of about 90%. So the download on the tail increases much more than the lift on the wing. This effect reduces the potential gain in lift, because the wing must carry the additional download. In practice, this effect reduces the potential benefit of the flaps to around half what the lift coefficient gain suggests it possible. Further, the increased tailplane size necessary adds to the cruise drag. You also need a larger vertical tail, to maintain directional stability - and increased dihedral to offset the spiral instability due to the larger vertical tail.

 

Secondly, the induced drag increases in proportion to the square of the lift coefficient - and there's an addition to the simple induced drag due to the variation of the drag from the flaps; this results in the minimum drag speed being increased; so the tendency towards speed instability ("on the back side of the drag curve") is markedly increased; extreme STOL aircraft are noted for tending to fall out of the air due to this; the Helio Courier etc were notorious for it.

 

Thirdly, the lateral stability is adversely affected because of the effect of the asymmetric slipstream interacting with a powerful flap system, in the crossed-control case required by the certification testing.

 

Fourthly, the behaviour of such aircraft in an aborted landing can be rather startling - in the case of the Helio Stallion, approaching lethal. Generally, the more powerful the flap system, the more powerful needs to be the engine - and the combination often produces a violent pitch-up tendency when power is applied to go around; there are limits set in FAR 23 to the stick force necessary to control this - they are there as a result of the Helio Stallion, which needed both hands on the wheel, so you could not re-trim the thing, because it lacked an electric trim capability. It's all "designing into a corner".

 

All these things can be catered for in the original design - the Twin Otter is a prime example - but there's a price for it. STOL was fashionable in the 1960s, with aircraft like the Helio Courier, turbo Beaver, Pilatus turbo-porter, Caribou, etc, which can show startling performance; however they did not make much market penetration except where special circumstances gave them an advantage. Ansett, for example, found to his cost that a Caribou, when operated under normal civil rules for balanced field length, could not lift as much freight as a DC-3, on identical engines & propellers. (It became known as the Caribou boo). In military use, which ignores the risk of engine failure on takeoff, the Caribou was remarkable. It either works, or you're dead.

 

Cessna found it advisable to remove the fourth flap notch from the 182, because the average pilot was barely able to contend with the side effects of the increased performance. The latest C172 has reduced flap deflection, so it's difficult to make a steep power-off approach in it. I can only assume this is a result of the cost of product liability insurance, but this is the way things are going.

...but back to the thread: A small increase in overall aeroplane Clmax, resulting in a (square root of small) decrease in stall speed, is almost certainly schievable; but the scope of modifications required to get far, are likely to be horrifically large. As Dafydd points out, at this edge of the envelope, the longitudinal stability, lateral stability, handling, stall behaviour, and balked landing behaviour are all in a contest to keep the minimum speed up...

The 2-seat Thrusters stall around 35~37kts, no flaps :o)

 

 

[Dafydd Llewellyn]

Yep; as far as the KR2 goes, my view is similar to that of the farmer who was asked for directions to Bungledoo - he scratched his head, looked up and down the road, and then said "If I wanted to get there, I wouldn't start from here."

 

 

[Head in the clouds]

....... *Where can I find a reference to that Bob? ......

 

NACA Technical Note 808 shows the two-dimensional lift coefficient of 23012 with a single 0.3c Fowler flap at 40 degrees deflection, as 3.3 compared to the naked section at 1.52; I think the same data is plotted out in Abbot & Von Doenhoff. This equates to a dCl of ~1.78; at KR-2 reynolds numbers and planform, the maximum practicable root Clmax (unflapped) will be about 1.4 If one assumes ~65% span flaps, with a bit of effective taper, then the fowler in question has the potential to just about double the Clmax (the inner 65% span carries about 80% of the load, giving an overall dCl of 1.78 x 0.8 = 1.42). This would give 1/1.42, or a ~30% reduction in Vs.NACA Report 723 shows 230XX in three thicknesses, with a moderately undeveloped 40% chord double-slooted flap pack. There is more information therein than meets the eye, as 23012 is a leading-edge staller, whereas 23030 is definitely a trailing-edge staller (23021 changes with Re, but may be considered TE staller for small a/c purposes). In that Report, 23012 had a max dCl of 1.9+; 23030 had ~2.67. As it's the upper rear wing section curvature that is most affected by the flap local suction (thus leading to greater circulation etc), the 23030 data should read across to a thinner aft-cambered TE stalling section, like USAF 35B. On the same basis as the Fowler considered above, the stall speed comes down to 59% of the flapless value - viola!.

A conservative aerodynamicist would regard this pasting of the 23030 dCl onto any old rear-loaded airfoil as an optomistic guess, as much because of the retrospectively known vagueness in the NACA correlations between excessive tunnel turbulence and effective reynolds number, as because of the paucity of data on short-bubble behaviour of 23030 (which should greatly affect the translatability of the data to 6-digit sections). It is of note that the 6-digit data published in Abbot & von Doenhoff predates the discovery - by Hoerner - of the VD tunnel's vast excess of turbulence; so the 6-digit Clmaxes given in A&vD are optomistic at least.

 

So - can we get a 40% reduction in Vs, or not? Well, NACA WR L310 - which is subsequent to the 40% double-slotted tests - shows that minor changes in the fore flap - aft flap relationship can give a dCl variation - on a 0.3c flap pack - of 0.56. As this research was based directly on the 0.4c double slotted flap pack, it is reasonable to assume that the 0.4c pack could be improved by as much as 0.3 (dCl) by similar tweaking; which is larger than the various tunnel errors discussed previously, and suggests that the previous generalisation about wide applicability of the 23030 data is defensible.

 

But wait, there's more! An ace lurks up the sleeve, in the form of some data - reported I think in Hoerner's "Fluid Dynamic Lift" - showing a remarkable benefit from a downturned flap TE on a slotted flap. Or perhaps not remarkable - if a plain flappette gives even a 50% improvement on flap nose suction (i.e. local dCl of ~0.6), then the knock-on effect should give a very similar magnitude of benefit to the whole wing / flap pack assembly - i.e. a 50% improvement in overall wing dCl.

 

But is the idea of an overall Clmax of >4 the ramblings of a madman? Schrenk's tests on blown fat airfoils got to a Clmax of well over 5 (from memory). If we regard the rear flappette as increasing the nose suction of the rear flap, which in turn increases the nose suction of the front flap, creating a boundary layer control equivalent to a wing with a suction slot at the TE, it becomes apparent that the overall wing circulation behaviour of a highly flapped wing should be very comparable to that of a 'forced' boundary layer wing.

 

It is of note that the airliners of today, which might be expected to be designed to later data than the published NACA work, are in fact using flap packs that are linear developments of the NACA work, and not very much different at all.

I had a look through your suggested references and if I'm interpreting them correctly the theory suggests a 40% stall speed might be possible. But, as I've said, the theoretical case is rarely achieved in the real world - I don't think even airliners actually achieve that with very developed and complex multiple slotted flaps (they get a little bit closer with slats in addition) but LSAs which are unlikely to have anything more than single slotted flaps won't get anywhere near that figure. Wouldn't you agree?

 

 

[Dafydd Llewellyn]

In the real World, the original Jabiru achieved a maximum lift coefficient for the whole aircraft (i.e., Cn max, not Cl max) 0f 2.2 with its single-slot flap at 33 degrees deflection. That's about as much flap as it can carry and still be able to climb in a baulked-landing situation. The practical reality is that a "balanced" LSA design with the normal sort of power and wing loadings, won't be able to use much more than this.

 

 

[Bob Llewellyn]

<reply to headintheclouds>

 

I presume we're both talking about a stall speed reduction, from the "clean" (no flaps / slats / blowing / suction / doovers) value, to 60% of that speed - for example, a reduction from 60 kts clean to 36 kts in landing configuration?

 

Without splitting further hairs, such an aeroplane could not be made certifiable under LSA; would not have much utility within the LSA weight limits; does not comply with the LSA "simple aeroplane" concept; and would require a lot of engineering, development, and money to achieve.

 

A solution to the very low stall speed aircraft exists, however - they're called "gyrocopters".

 

The painful reality Dafydd keeps alluding to, is that any certifiable aeroplane optimised for ton-miles per dollar-hour will find it's flap power to be limited by pitch control authority. There is one known "improvement" to be had (in this respect) over the single-slotted hinged flap, which is the Fowler; by the NACA data as plotted in A&vD, a lift coefficient of up to 2.8 (2D) can be had with about half the pitching moment increment of simple hinged flaps.

 

The Fowler has the additional benefit that, when the deflection is increased, the Clmax decreases slightly, and the drag goes up markedly. This means that one can use them as landing flaps in the high-drag position, and - if a balked landing becomes necessary - dumping a notch of flap both reduces drag and increases the available lift. I suspect that this is now considered too complex for a "pilot of average abilities"...

 

The downside of the Fowler is the mechanism required to push the flap 0.3c (or whatever) rearwards... especially as the cable interconnect used by Cessna is now frowned upon by design standards*.

 

*Certifying Authorities actually; as the flap interconnect is also the flap drive, and is a complex system with a bunch of single failure modes resulting in catastrophic accident conditions...Since I have yet to see any recreational aeroplane with a properly-shaped flap shroud - despite the data having been in the public domain for the last 80 years - I have to conclude that most LSA designers aren't serious about keeping stall speeds down.

 

 

[Dafydd Llewellyn]

That's pretty much the reality of it. However, practical experience (on the Seabird Seeker) shows that a suitable configuration of VGs can increase the maximum lift coefficient by about 13% - which would take the Jabiru value from 2.2 to just under 2.5 - which would result in a stall speed reduction of about 2.5 knots - PROVIDED there was sufficient elevator power to allow the necessary increase in the angle of attack. Also, PROVIDED the design of the VG layout is correct - and most of them that I've seen are not - this can be achieved with very good stall handling behaviour. What Seabird did to achieve that is their proprietory information. It's a hazardous area for experimentation.

 

 

[rgmwa]

Since I have yet to see any recreational aeroplane with a properly-shaped flap shroud - despite the data having been in the public domain for the last 80 years - I have to conclude that most LSA designers aren't serious about keeping stall speeds down.

I assume you mean where the back of the wing has an `S' shape profile to direct the air from the bottom of the wing up through the slot and back over the extended flap?

 

rgmwa

 

 

[Bob Llewellyn]

I assume you mean where the back of the wing has an `S' shape profile to direct the air from the bottom of the wing up through the slot and back over the extended flap?rgmwa

yup, that's it. NACA Report 664 shows in detail the relationship between the tested flap nose profiles and the shroud / lip at several deflections; it's of note that the flap first drops down, then shifts back and up as the angle of deflection increases. This maintains sufficient energy in the jet of air impinging on the flap nose (and flowing over the upper surface) to avoid separation. The motion is quite complex, and very minor geometrical errors will lose a lot of the targetted performance. Most light aeroplanes use a single hinge to move the flap, which can only be optimum for a single deflection; but most light aeroplanes do not have a shroud OR lip configuration within half a mile of correct at ANY deflection.

 

 

[spacesailor]

AND the last, Soar like an eagle, Land like a feather,

 

with a big Parachute, like in the Blue mountains nsw

 

spacesailor

 

 

[Bob Llewellyn]

AND the last, Soar like an eagle, Land like a feather,with a big Parachute, like in the Blue mountains nsw

 

spacesailor

I've seen some pretty thumpy eagle landings...