Vertical stablizer
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What specifically do you want to know?
Can you explain your theory (well, your friend's actually) of boundary layer thickening at the tips increasing the wing stalling angle on a swept wing?
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Can you explain your theory (well, your friend's actually) of boundary layer thickening at the tips increasing the wing stalling angle on a swept wing?
Now, many early jet transport types had vortex generators for boundry layer control, specifically early models of the B707, on the wing in front of the inboard aileron and on the underside of the horizontal tailplane, just forward of the elevators....to energise the boundry layer, to facilitate proper control surface effectiveness.
What do your texts say about boundry layer 'thickening at the wing tips?'
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Well, old chap - in posts #9 and #19 you said that spanwise flow was the reason for the higher stalling angles. I said (in #14) that spanwise flow was, in my understanding, a 'red herring' in terms of answering the OP's question, since when you have been all around the houses with 'expert friends' in Arizona who are now 'all at sea', but you still have not addressed the question! We are still seeking an explanation to try to help the OP, who did not ask about boundary layer thickening at the tips, but did ask about the increase in stalling angle.
Is there anyone who actually knows?
If I remember correctly, DPD only states the increase as a fact with no explanation.
Is there anyone who actually knows?
If I remember correctly, DPD only states the increase as a fact with no explanation.
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If I remember correctly, DPD only states the increase as a fact with no explanation.
I will repeat...NO one single individual influenced the design of jet transport aircraft the way DP Davies...did.
I first read his first edition, circa 1968, just prior to flying the B707.
That particular design was very unforgiving...with the early types.
Later ones...very nice.
I would certainly like to think that Captain Davies had a hand in all this...and he most certainly did.
Especially...with the rudder power boost system (it already had balance panels...IE: servo tabs) and a parallel yaw damper.
As, the parallel yaw damper had to be switched OFF, for takeoff and landing.
The airplane was a real handful, in gusty winds, without the yaw damper engaged.
I kid you not!
Later, series yaw damper...superb.
And, you could split the spoilers for pitch control...very effective, if needed.
Only once did I need to split the spoilers...into the old Taipei airport, long ago.
IE; jammed stab.
Last edited by 411A; 19th Oct 2009 at 19:22.
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Explanation sounds OK to me, but then I'll believe anything.
ASK DJ Aerotech Question
Sweep affects the lift-curve slope. The lift curve slope ("dCl/d_alpha" in engineering jargon) is nothing more than how much the lift coefficient ("Cl") changes for a given change in the angle of attack ("alpha"). For most airfoils, the plot of Cl vs. alpha is nearly a straight line except for some rounding off near the stall, so the dCl/d_alpha is nearly constant for most of the operating range. A typical value is about 0.1 (i.e.: the lift coefficient increases by about 0.1 for each 1 degree increase in angle of attack).
Sweeping a wing changes its lift curve slope, so the new dCl/d-alpha is equal to the unswept dCl/d-alpha times the cosine of the sweep angle. For example, if the unswept dCl/d-alpha was 0.1, a sweep angle of 30 degrees would result in a dCl/d-alpha of only 0.087 . This also explains why swept-wing airplanes can get to much higher angles of attack before stalling. The airfoil stalls at pretty much the same lift coefficient regardless of what the sweep angle is (until we get to extreme sweep angles that result in significant amounts of vortex lift), but the decrease in dCl/d-alpha due to the sweep means that we need to have a higher angle of attack to reach the stalling Cl.
ASK DJ Aerotech Question
Sweep affects the lift-curve slope. The lift curve slope ("dCl/d_alpha" in engineering jargon) is nothing more than how much the lift coefficient ("Cl") changes for a given change in the angle of attack ("alpha"). For most airfoils, the plot of Cl vs. alpha is nearly a straight line except for some rounding off near the stall, so the dCl/d_alpha is nearly constant for most of the operating range. A typical value is about 0.1 (i.e.: the lift coefficient increases by about 0.1 for each 1 degree increase in angle of attack).
Sweeping a wing changes its lift curve slope, so the new dCl/d-alpha is equal to the unswept dCl/d-alpha times the cosine of the sweep angle. For example, if the unswept dCl/d-alpha was 0.1, a sweep angle of 30 degrees would result in a dCl/d-alpha of only 0.087 . This also explains why swept-wing airplanes can get to much higher angles of attack before stalling. The airfoil stalls at pretty much the same lift coefficient regardless of what the sweep angle is (until we get to extreme sweep angles that result in significant amounts of vortex lift), but the decrease in dCl/d-alpha due to the sweep means that we need to have a higher angle of attack to reach the stalling Cl.
simple at low altitudes dynamic pressure effects predominate, at high altitudes compresibilty effects dominate along with the deleterious effects of reduced aero dynamic damping--- sweep delays onset of this whol mess awhile
that's why an ejection at the speed /altitude limits of an SR 71 wont rip the pilot to smithereens but at low altitude/high dymanic pressure conditions it must be quite a ride
nowadays that they don't make too many turbojets let's for get the coffin corner stuff mostly
PA
that's why an ejection at the speed /altitude limits of an SR 71 wont rip the pilot to smithereens but at low altitude/high dymanic pressure conditions it must be quite a ride
nowadays that they don't make too many turbojets let's for get the coffin corner stuff mostly
PA
Last edited by Pugilistic Animus; 19th Oct 2009 at 23:55.
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Originally Posted by BA
For example, if the unswept dCl/d-alpha was 0.1, a sweep angle of 30 degrees would result in a dCl/d-alpha of only 0.087 . This also explains why swept-wing airplanes can get to much higher angles of attack before stalling.
Here's a gentle puzzle too - since sweep is measured at mean chord, what is the theoretical maximun sweep angle?
BOAC what I'm trying to actyually ask is:
against what limitation do you mean when you refere to 'maximal sweep' ?
and some of it will depend on empirical airfoil characteristics and the overall section characterisics in terms of cl/CL
PA
against what limitation do you mean when you refere to 'maximal sweep' ?
and some of it will depend on empirical airfoil characteristics and the overall section characterisics in terms of cl/CL
PA
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Hope this contributes and sheds some light.
Effect of Sweepback on Lift
If a straight wing is changed to a swept planform, with similar parameters of area, aspect ratio, taper, section and washout, the CLmax is reduced. This is due to premature flow separation from the upper surface at the wing tips. For a sweep angle of 45º, the approximate reduction in CLmax is around 30%. Fig 3 shows typical CL curves for a straight wing, a simple swept-back wing, and a tailless delta wing of the same low aspect ratio.
Fig 3 Effect of Planform on CLmax
A swept wing presents less camber and a greater fineness ratio to the airflow. However, the reasons for the lowering of the CL slope is more readily apparent from an examination of Figs 4 and 5. From Fig 4 it can be seen that the velocity V can be divided into two components, V1 parallel to the leading edge which has no effect on the lift, and V2 normal to the leading edge which does affect the lift and is equal to V cos Λ. Therefore, all other factors being equal, the CL of a swept wing is reduced in the ratio of the cosine of the sweep angle.
An alternative explanation is that the component parallel to the leading edge produces no lift. Only the component normal to the leading edge is considered to be producing lift. As this component is always less than the free stream flow at all angles of sweep, a swept wing will always produce less lift than a straight wing.
Fig 4 Flow Velocities on a Swept Wing
Fig 5 shows that an increase in fuselage angle of attack Δα will only produce an increase in the angle of attack Δα cos Λ in the plane perpendicular to the wing quarter chord line. Since we have already said that it is airflow in the latter plane which effects CL, the full increment of lift expected from the Da change is reduced to that of a Da cos L change.
Fig 5 Effect of change in Angle of Attack
Considering Fig 3, the stall occurs on all three wings at angles of attack considerably greater than those of wings of medium and high aspect ratios. On all aircraft it is desirable that the landing speed should be close to the lowest possible speed at which the aircraft can fly; to achieve this desirable minimum the wing must be at the angle of attack corresponding to the CLmax.
On all wings of very low aspect ratio, and particularly on those with a swept-back planform, the angles of attack giving the highest lift coefficients cannot be used for landing. This is because, as explained later, swept-back planforms have some undesirable characteristics near the stall and because the exaggerated nose-up attitude of the aircraft necessitates, among other things, excessively long and heavy undercarriages. The maximum angle at which an aircraft can touch down without recourse to such measures is about 15º, and the angle of attack at touch-down will therefore have to be something of this order. Fig 3 shows that the CL corresponding to this angle of attack is lower than the CLmax for each wing.
Compared with the maximum usable lift coefficient available for landing aircraft with unswept wings, those of the swept and delta wings are much lower, necessitating higher landing speeds for a given wing loading. It is now apparent that, to obtain a common minimum landing speed at a stated weight, an unswept wing needs a smaller area than either of the swept planforms. The simple swept wing needs a greater area, and so a lower wing loading, in order that the reduced CL can support the weight at the required speed. The tailless delta wing needs still more area, and so a still lower wing loading, to land at the required speed. Fig 6 shows typical planforms for the three types of wing under consideration, with the areas adjusted to give the same stalling speed. The much larger area of the delta wing is evident.
Figure 6 Comparative Planforms
A NASA of early1950's vintage paper on low speed swept wing performance here.
http://ntrs.nasa.gov/archive/nasa/ca...1993086214.pdf
Effect of Sweepback on Lift
If a straight wing is changed to a swept planform, with similar parameters of area, aspect ratio, taper, section and washout, the CLmax is reduced. This is due to premature flow separation from the upper surface at the wing tips. For a sweep angle of 45º, the approximate reduction in CLmax is around 30%. Fig 3 shows typical CL curves for a straight wing, a simple swept-back wing, and a tailless delta wing of the same low aspect ratio.
Fig 3 Effect of Planform on CLmax
A swept wing presents less camber and a greater fineness ratio to the airflow. However, the reasons for the lowering of the CL slope is more readily apparent from an examination of Figs 4 and 5. From Fig 4 it can be seen that the velocity V can be divided into two components, V1 parallel to the leading edge which has no effect on the lift, and V2 normal to the leading edge which does affect the lift and is equal to V cos Λ. Therefore, all other factors being equal, the CL of a swept wing is reduced in the ratio of the cosine of the sweep angle.
An alternative explanation is that the component parallel to the leading edge produces no lift. Only the component normal to the leading edge is considered to be producing lift. As this component is always less than the free stream flow at all angles of sweep, a swept wing will always produce less lift than a straight wing.
Fig 4 Flow Velocities on a Swept Wing
Fig 5 shows that an increase in fuselage angle of attack Δα will only produce an increase in the angle of attack Δα cos Λ in the plane perpendicular to the wing quarter chord line. Since we have already said that it is airflow in the latter plane which effects CL, the full increment of lift expected from the Da change is reduced to that of a Da cos L change.
Fig 5 Effect of change in Angle of Attack
Considering Fig 3, the stall occurs on all three wings at angles of attack considerably greater than those of wings of medium and high aspect ratios. On all aircraft it is desirable that the landing speed should be close to the lowest possible speed at which the aircraft can fly; to achieve this desirable minimum the wing must be at the angle of attack corresponding to the CLmax.
On all wings of very low aspect ratio, and particularly on those with a swept-back planform, the angles of attack giving the highest lift coefficients cannot be used for landing. This is because, as explained later, swept-back planforms have some undesirable characteristics near the stall and because the exaggerated nose-up attitude of the aircraft necessitates, among other things, excessively long and heavy undercarriages. The maximum angle at which an aircraft can touch down without recourse to such measures is about 15º, and the angle of attack at touch-down will therefore have to be something of this order. Fig 3 shows that the CL corresponding to this angle of attack is lower than the CLmax for each wing.
Compared with the maximum usable lift coefficient available for landing aircraft with unswept wings, those of the swept and delta wings are much lower, necessitating higher landing speeds for a given wing loading. It is now apparent that, to obtain a common minimum landing speed at a stated weight, an unswept wing needs a smaller area than either of the swept planforms. The simple swept wing needs a greater area, and so a lower wing loading, in order that the reduced CL can support the weight at the required speed. The tailless delta wing needs still more area, and so a still lower wing loading, to land at the required speed. Fig 6 shows typical planforms for the three types of wing under consideration, with the areas adjusted to give the same stalling speed. The much larger area of the delta wing is evident.
Figure 6 Comparative Planforms
A NASA of early1950's vintage paper on low speed swept wing performance here.
http://ntrs.nasa.gov/archive/nasa/ca...1993086214.pdf
it great stuff Brian Abraham I have some great publications on this subject also derived from NASA/NACA data, but have not seen the originals I love these historical items in engineering
PA
PA
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Brian - that is where I got to. Incidentally, do you know what 'A=2' means in Fig3? That fig is a little confusing too since it shows the left as 'no sweep' and then talks later about the 'the stall occurs on all three wings etc etc'??
I have asked all the Bsc Aero-eng's of my time and none have actually come up with an explanation of "Considering Fig 3, the stall occurs on all three wings at angles of attack considerably greater than those of wings of medium and high aspect ratios.". Do you have any reference to why? To my mind, ignoring CL and the vector analysis of the flow (which is useful for the 'mathematics'), the air ACTUALLY flows more or less from front to back in the line of flight (with a small touch of outflow, of course) and I am puzzled as to why the swept wing is reaching higher angles before stalling. It does flow over a 'thinner' section due to the sweep, but is that enough to raise apha crit?
I see that the NASA paper does indeed, as I suggested earlier, delve into separated vortices and I still wonder if this is the reason for the delayed 'stall'?
I have asked all the Bsc Aero-eng's of my time and none have actually come up with an explanation of "Considering Fig 3, the stall occurs on all three wings at angles of attack considerably greater than those of wings of medium and high aspect ratios.". Do you have any reference to why? To my mind, ignoring CL and the vector analysis of the flow (which is useful for the 'mathematics'), the air ACTUALLY flows more or less from front to back in the line of flight (with a small touch of outflow, of course) and I am puzzled as to why the swept wing is reaching higher angles before stalling. It does flow over a 'thinner' section due to the sweep, but is that enough to raise apha crit?
I see that the NASA paper does indeed, as I suggested earlier, delve into separated vortices and I still wonder if this is the reason for the delayed 'stall'?
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G'day BOAC, The A=2 is Aspect Ratio. Forgot to say the previous and this AP3456.
A stall occurs when the effective angle of attack reaches the critical angle. Induced downwash reduces the effective angle of attack of a wing. Since induced drag is inversely proportional to aspect ratio it follows that a low aspect ratio wing will have high induced drag, high induced downwash and a reduced effective angle of attack. The low aspect ratio wing therefore has a higher stalling angle of attack than a wing of high aspect ratio.
The reduced effective angle of attack of very low aspect ratio wings can delay the stall considerably. Some delta wings have no measurable stalling angle up to 40º or more inclination to the flight path. At this sort of angle the drag is so high that the flight path is usually inclined downwards at a steep angle to the horizontal. Apart from a rapid rate of descent, and possible loss of stability and control, such aircraft may have a shallow attitude to the horizon and this can be deceptive. The condition is called the super stall or deepstall, although the wing may be far from a true stall and still be generating appreciable lift.
When a wing is swept back, the boundary layer tends to change direction and flow towards the tips. This outward drift is caused by the boundary layer encountering an adverse pressure gradient and flowing obliquely to it over the rear of the wing.
The pressure distribution on a swept wing is shown by isobars in Fig 8. The velocity of the flow has been shown by two components, one at right angles and one parallel to the isobars. Initially, when the boundary layer flows rearwards from the leading edge it moves towards a favourable pressure gradient, ie towards an area of lower pressure. Once past the lowest pressure however, the component at right angles to the isobars encounters an adverse pressure gradient and is reduced. The component parallel to the isobars is unaffected, thus the result is that the actual velocity is reduced (as it is over an unswept wing) and also directed outwards towards the tips.
The direction of the flow continues to be changed until the component at right angles to the isobars is reduced to zero, whilst the parallel component, because of friction, is also slightly reduced. This results in a “pool” of slow moving air collecting at the tips.
The spanwise drift sets up a tendency towards tip stalling, since it thickens the boundary layer over the outer parts of the wing and makes it more susceptible to separation, bringing with it a sudden reduction in CLmax over the wing tips.
At the same time as the boundary layer is flowing towards the tips, at high angles of attack, the airflow is separating along the leading edge. Over the inboard section it re-attaches behind a short “separation bubble”, but on the outboard section it re-attaches only at the trailing edge or fails to attach at all. The separated flow at the tips combines with the normal wing tip vortices to form a large vortex (the ram’s horn vortex). The factors which combine to form this vortex are:
a. Leading edge separation.
b. The flow around the wing tips.
c. The spanwise flow of the boundary layer.
These factors are illustrated in Fig 9, and the sequence of the vortex development and its effect on the airflow over the wing is shown in Fig 10. From Fig 10 it can be seen that the ram’s horn vortex has its origin on the leading edge, possibly as far inboard as the wing root. The effect of the vortex on the air above it (the external flow) is to draw the latter down and behind the wing, deflecting it towards the fuselage (Fig 11).
The spanwise flow of the boundary layer increases as angle of attack is increased. This causes the vortex to become detached from the leading edge closer inboard (see Fig 12). As a result, outboard ailerons suffer a marked decrease in response with increasing angle of attack. This, in turn, means that comparatively large aileron movements are necessary to manoeuvre the aircraft at low speeds; the aircraft response may be correspondingly sluggish. This effect may be countered by limiting the inboard encroachment of the vortex as described below, or by moving the ailerons inboard. Another possible solution is the use of an all-moving wing tip.
Most of the methods used to alleviate the tip stall aim either at maintaining a thin and therefore strong boundary layer, or re-energizing the weakened boundary layer:
a. Boundary Layer Fences. Used originally to restrict the boundary layer out-flow, fences also check the spanwise growth of the separation bubble along the leading edge.
b. Leading Edge Slots. These have the effect of re-energizing the boundary layer.
c. Boundary Layer Suction. Suitably placed suction points draw off the weakened layer; a new high-energy layer is then drawn down to take its place.
d. Boundary Layer Blowing. High velocity air is injected into the boundary layer to increase its energy.
e. Vortex Generators. These re-energize the boundary layer by making it more turbulent. The increased turbulence results in high-energy air in layers immediately above the retarded layer being mixed in and so re-energizing the layer as a whole. Vortex generators are most commonly fitted ahead of control surfaces to increase their effect by speeding up and strengthening the boundary layer. Vortex generators also markedly reduce shock-induced boundary layer separation, and reduce the effects of the upper surface shockwave.
f. Leading Edge Extension. Also known as a “sawtooth” leading edge, the extended leading edge is a common method used to avoid the worst effects of tip stalling. The effect of the extension is to cut down the growth of the main vortex. A further smaller vortex, starting from the tip of the extension, affects a much smaller proportion of the tip area and in lying across the wing, behind the tip of the extension, it has the effect of restricting the outward flow of the boundary layer. In this way the severity of the tip stall is reduced and with it the pitch-up tendency. Further effects of the leading edge extension are:
(1) The t/c ratio of the tip area is reduced, with consequent benefits to the critical Mach number.
(2) The Centre of Pressure (CP) of the extended portion of the wing lies ahead of what would be the CP position if no extension were fitted. The mean CP position for the whole wing is therefore further forward and, when the tip eventually stalls, the forward shift in CP is less marked, thus reducing the magnitude of the nose-up movement.
g. Leading Edge Notch. The notched leading edge has the same effect as the extended leading edge insofar as it causes a similar vortex formation thereby reducing the magnitude of the vortex over the tip area and with it the tip stall. Pitch-up tendencies are therefore reduced. The leading edge notch can be used in conjunction with extended leading edge, the effect being to intensify the inboard vortex behind the devices to create a stronger restraining effect on boundary layer out-flow. The choice whether to use either or both of these devices lies with the designer and depends on the flight characteristics of the aircraft.
Aspect Ratio and Stalling Angle
A stall occurs when the effective angle of attack reaches the critical angle. Induced downwash reduces the effective angle of attack of a wing. Since induced drag is inversely proportional to aspect ratio it follows that a low aspect ratio wing will have high induced drag, high induced downwash and a reduced effective angle of attack. The low aspect ratio wing therefore has a higher stalling angle of attack than a wing of high aspect ratio.
The reduced effective angle of attack of very low aspect ratio wings can delay the stall considerably. Some delta wings have no measurable stalling angle up to 40º or more inclination to the flight path. At this sort of angle the drag is so high that the flight path is usually inclined downwards at a steep angle to the horizontal. Apart from a rapid rate of descent, and possible loss of stability and control, such aircraft may have a shallow attitude to the horizon and this can be deceptive. The condition is called the super stall or deepstall, although the wing may be far from a true stall and still be generating appreciable lift.
Effect of Sweepback on Stalling
When a wing is swept back, the boundary layer tends to change direction and flow towards the tips. This outward drift is caused by the boundary layer encountering an adverse pressure gradient and flowing obliquely to it over the rear of the wing.
The pressure distribution on a swept wing is shown by isobars in Fig 8. The velocity of the flow has been shown by two components, one at right angles and one parallel to the isobars. Initially, when the boundary layer flows rearwards from the leading edge it moves towards a favourable pressure gradient, ie towards an area of lower pressure. Once past the lowest pressure however, the component at right angles to the isobars encounters an adverse pressure gradient and is reduced. The component parallel to the isobars is unaffected, thus the result is that the actual velocity is reduced (as it is over an unswept wing) and also directed outwards towards the tips.
Fig 8 Outflow of Boundary Layer
The direction of the flow continues to be changed until the component at right angles to the isobars is reduced to zero, whilst the parallel component, because of friction, is also slightly reduced. This results in a “pool” of slow moving air collecting at the tips.
The spanwise drift sets up a tendency towards tip stalling, since it thickens the boundary layer over the outer parts of the wing and makes it more susceptible to separation, bringing with it a sudden reduction in CLmax over the wing tips.
At the same time as the boundary layer is flowing towards the tips, at high angles of attack, the airflow is separating along the leading edge. Over the inboard section it re-attaches behind a short “separation bubble”, but on the outboard section it re-attaches only at the trailing edge or fails to attach at all. The separated flow at the tips combines with the normal wing tip vortices to form a large vortex (the ram’s horn vortex). The factors which combine to form this vortex are:
a. Leading edge separation.
b. The flow around the wing tips.
c. The spanwise flow of the boundary layer.
These factors are illustrated in Fig 9, and the sequence of the vortex development and its effect on the airflow over the wing is shown in Fig 10. From Fig 10 it can be seen that the ram’s horn vortex has its origin on the leading edge, possibly as far inboard as the wing root. The effect of the vortex on the air above it (the external flow) is to draw the latter down and behind the wing, deflecting it towards the fuselage (Fig 11).
Fig 9 Vortex Development
Fig 10 Formation of Ram’s Horn Vortex
Fig 11 lnfluence on External Flow
The spanwise flow of the boundary layer increases as angle of attack is increased. This causes the vortex to become detached from the leading edge closer inboard (see Fig 12). As a result, outboard ailerons suffer a marked decrease in response with increasing angle of attack. This, in turn, means that comparatively large aileron movements are necessary to manoeuvre the aircraft at low speeds; the aircraft response may be correspondingly sluggish. This effect may be countered by limiting the inboard encroachment of the vortex as described below, or by moving the ailerons inboard. Another possible solution is the use of an all-moving wing tip.
Fig 12 Shift of Ram’s Horn Vortex
Alleviating the Tip Stall
Most of the methods used to alleviate the tip stall aim either at maintaining a thin and therefore strong boundary layer, or re-energizing the weakened boundary layer:
a. Boundary Layer Fences. Used originally to restrict the boundary layer out-flow, fences also check the spanwise growth of the separation bubble along the leading edge.
b. Leading Edge Slots. These have the effect of re-energizing the boundary layer.
c. Boundary Layer Suction. Suitably placed suction points draw off the weakened layer; a new high-energy layer is then drawn down to take its place.
d. Boundary Layer Blowing. High velocity air is injected into the boundary layer to increase its energy.
e. Vortex Generators. These re-energize the boundary layer by making it more turbulent. The increased turbulence results in high-energy air in layers immediately above the retarded layer being mixed in and so re-energizing the layer as a whole. Vortex generators are most commonly fitted ahead of control surfaces to increase their effect by speeding up and strengthening the boundary layer. Vortex generators also markedly reduce shock-induced boundary layer separation, and reduce the effects of the upper surface shockwave.
f. Leading Edge Extension. Also known as a “sawtooth” leading edge, the extended leading edge is a common method used to avoid the worst effects of tip stalling. The effect of the extension is to cut down the growth of the main vortex. A further smaller vortex, starting from the tip of the extension, affects a much smaller proportion of the tip area and in lying across the wing, behind the tip of the extension, it has the effect of restricting the outward flow of the boundary layer. In this way the severity of the tip stall is reduced and with it the pitch-up tendency. Further effects of the leading edge extension are:
(1) The t/c ratio of the tip area is reduced, with consequent benefits to the critical Mach number.
(2) The Centre of Pressure (CP) of the extended portion of the wing lies ahead of what would be the CP position if no extension were fitted. The mean CP position for the whole wing is therefore further forward and, when the tip eventually stalls, the forward shift in CP is less marked, thus reducing the magnitude of the nose-up movement.
g. Leading Edge Notch. The notched leading edge has the same effect as the extended leading edge insofar as it causes a similar vortex formation thereby reducing the magnitude of the vortex over the tip area and with it the tip stall. Pitch-up tendencies are therefore reduced. The leading edge notch can be used in conjunction with extended leading edge, the effect being to intensify the inboard vortex behind the devices to create a stronger restraining effect on boundary layer out-flow. The choice whether to use either or both of these devices lies with the designer and depends on the flight characteristics of the aircraft.
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Just wish I'd kept my AP3456 - a brilliant publication.
It went out years ago (with my hat)
So, is the key that the Aspect Ratio of a swept wing is lower than that of a straight and are we indeed looking at vortex separation as the key to the higher alpha crit? Something has to be keeping the flow attached at those angles.
As you/it says, take AR to 'ridiculous' extremes as in the HP115 and Concorde research stuff and we are looking at around 40 dgerees. They are mainly vortex lifters, so is there a gradual transition of lift generation from 'traditional Bernoulli' to vortex as sweep increases? Is there indeed any evidence of a 'mini' vortex on a straight wing?
It went out years ago (with my hat)So, is the key that the Aspect Ratio of a swept wing is lower than that of a straight and are we indeed looking at vortex separation as the key to the higher alpha crit? Something has to be keeping the flow attached at those angles.
As you/it says, take AR to 'ridiculous' extremes as in the HP115 and Concorde research stuff and we are looking at around 40 dgerees. They are mainly vortex lifters, so is there a gradual transition of lift generation from 'traditional Bernoulli' to vortex as sweep increases? Is there indeed any evidence of a 'mini' vortex on a straight wing?
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Both Brian A and I are dusting off both our brains and our books to try and answer the OP's question.
It is extremely interesting that in another thread PMorten has unearthed a NASA paper on the subject which, although primarily involved in researching the methods of stopping the inherent instability of a traditionally swept wing at increasing alpha due to tip stalling, does disclose a significantly increasing l edge vortex effect which delays the stall. Best pics are at Fig 13 on page 33. I still pin my answer on that.
It is extremely interesting that in another thread PMorten has unearthed a NASA paper on the subject which, although primarily involved in researching the methods of stopping the inherent instability of a traditionally swept wing at increasing alpha due to tip stalling, does disclose a significantly increasing l edge vortex effect which delays the stall. Best pics are at Fig 13 on page 33. I still pin my answer on that.