P-40 turn rate/Flight model check

[JtD]

Please remember that the wing of the P-40 is inclined by 2°, so a stalling angle of attack of 14° gives a wing angle of attack of 16° (provided the plane is modelled with the same reference system like the original). Other aircraft have lower, variable or even 0° of wing inclination and can therefore achieve higher angles of attack without actually putting the wing chord at higher angles of attack.

[unreasonable]

Please remember that the wing of the P-40 is inclined by 2°, so a stalling angle of attack of 14° gives a wing angle of attack of 16° (provided the plane is modelled with the same reference system like the original). Other aircraft have lower, variable or even 0° of wing inclination and can therefore achieve higher angles of attack without actually putting the wing chord at higher angles of attack.

 

I was assuming the Tech Specs data was directly comparable - ie was the wing angle. Perhaps I had better ask in Qs for Developers unless someone already knows how the Tech Spec AoA number is defined.

 

16 degrees would indeed line up with the charts for similar airfoils.

[Farky]

Please remember that the wing of the P-40 is inclined by 2°, so a stalling angle of attack of 14° gives a wing angle of attack of 16° (provided the plane is modelled with the same reference system like the original). Other aircraft have lower, variable or even 0° of wing inclination and can therefore achieve higher angles of attack without actually putting the wing chord at higher angles of attack.

 

Small correction - wing of the P-40 is inclined by 1°, not 2°.

 

 

[Cpt_Branko]

Generally, wing has a smaller Clmax at a higher critical AoA than 2d airfoil profile. I just don't see 14 degrees as very plausible.

 

Anyway, developers can easily get the data, given it is in British archives. I'm really not prepared to shell out a hundred euros for it, but if you're selling virtual airplanes at 20$ a pop, well

[JtD]

Small correction - wing of the P-40 is inclined by 1°, not 2°.

I can't check my sources right now, and said that from the top of my head. Quite possible I'm wrong. Is the 1° backed up by multiple sources or did you just check the one you posted?

[19//Moach]

it is still possible, given how they overlooked the washout angle of the fw190, (the spanwise twist that gives the tip lower AoA than the root) that this (and similar details) could have been missed on the p40 as well -- likewise, the p40 has one or two degrees of washout on its wings

 

 

it does FEEL like when it stalls it more or less "simply goes" - instead of the escalating buffeting you'd expect from having the root stall before the tips - in a way, I always felt it to have something in common with the 190 (before 2.007), in how it performs near and through stalls

 

 

the correlation hints at a curiously possible relation between the two

 

but that's just my personal opinion - yet, since the 190 had such items missed, it'd make sense to at least double-check if any similar "oopsidaisies" can be found on the p40... just in case

 

 

 

anyways - here's a really cool source of extremely fine detail on the p40: http://www.p40warhawk.com/Models/Technical/Technical.htm

Edited February 1, 2017 by 19.GIAP//Moach

[Farky]

I can't check my sources right now, and said that from the top of my head. Quite possible I'm wrong. Is the 1° backed up by multiple sources or did you just check the one you posted?

 

Backed up by multiple sources.

[JtD]

I found 1.3° in my sources.

 

But, on another issue, something I wanted to look up. As you may know, the lift coefficient is generally connected to a reference area, which is calculated using wing span and average chord. It includes the fuselage, which has little impact on lift, and excludes tail area, which does have an impact.

 

Looking at the stalling condition of flight, meaning a high angle of attack, I wanted to know how the reference area compares to the actual airfoil area significantly contributing to lift. Since the elevator angle is usually upwards and, depending on aircraft trim, does not contribute to lift (could even produce a downforce), I've decided to ignore the elevators. I'm comparing reference area to wing area (without fuselage) plus area of horizontal stabilizer. I took the figures mostly from drawings, unless I found the stated explicitly. Comparing P-40 with Fw190 and Yak-1, it looks like that:

 

P-40 - 21.92 -> 22.4

Fw190 - 18.3 -> 17.6

Yak-1 - 17.15 -> 16.2

 

The aircraft lose about 2.3-2.5m² of reference area due to the fuselage, but the horizontal stabilizer sizes are considerably different. Only on the P-40 it is larger than the area lost to the fuselage, so the P-40 does not only lose a smaller percentage of the reference area to the fuselage, but is the only aircraft to have lifting surfaces larger than the reference area.

Bottom line, if the airfoils all came with the same 1.4 lift coefficient over all the surfaces, the coefficients for the reference area would read:

P-40: 1.43

Fw190: 1.35

Yak-1: 1.32

Meaning that just by plane geometry, the lift coefficient of the P-40 would be noticeably larger than those of the Fw190 and Yak-1. For our purposes that is, in game lift calculation is more complex.

 

Given that I'm sceptical towards anything above 1.4, at least for the P-40 I can raise that limit to 1.5 now.

Edited February 20, 2017 by JtD

[BlitzPig_EL]

So, how would this upward change in CLmax for the P40 E change the behavior of the aircraft in the sim?

 

Better turn even with it's low power to weight ratio?

 

Better maneuverability in general?

 

Lower stall speed clean and "dirty", which would lead to lower landing and takeoff speeds, which would be more in line with real world data?

Edited February 19, 2017 by BlitzPig_EL

[JtD]

Well, right now dev data says best turn performance of the P-40 is 24.3s@270km/h@sea level. This gives a lift coefficient of 1.1 and a turning radius of 290m. The Fw190 and the Yak-1's both utilize a lift coefficient of ~1.3, the Fw190 with a turning radius of 290m, the Yak-1's with 230m. If the P-40E lift coefficient was raised to ~1.4 in such a turn, without change in aerodynamic efficiency (basically saying the aircraft stays the same), turn time would go down to 23s, speed to 230km/h and turn radius to 230m. This would turn the P-40 into one of the best low speed turners in game, not with the best sustained time, but with one of the lowest turning speeds and smallest turning radius'. It wouldn't have any trouble turning inside of a 190, and could even hope to turn fight a 109 and live to tell the tale.

 

And yes, stall speeds would go down, so the differences between real world data and game performance would be smaller, however, there'd still be a considerable gap. This gap mind you is not necessarily an error, as we haven't managed to pinpoint the reason for it.

[BlitzPig_EL]

Thank you.

 

Seems this would put the P40 more in line with the things pilots said about it.

 

At least it would give you some options in combat other than just diving at an enemy then running away.

[DD_Arthur]

Seems this would put the P40 more in line with the things pilots said about it.

 

At least it would give you some options in combat other than just diving at an enemy then running away.

 

Amen.   Perhaps we could even have a good look at engine limits too.

[unreasonable]

I found 1.3° in my sources.

 

But, on another issue, something I wanted to look up. As you may know, the lift coefficient is generally connected to a reference area, which is calculated using wing span and average chord. It includes the fuselage, which has little impact on lift, and excludes tail area, which does have an impact.

 

Looking at the stalling condition of flight, meaning a high angle of attack, I wanted to know how the reference area compares to the actual airfoil area significantly contributing to lift. Since the elevator angle is usually upwards and, depending on aircraft trim, does not contribute to lift (could even produce a downforce), I've decided to ignore the elevators. I'm comparing reference area to wing area (without fuselage) plus area of horizontal stabilizer. I took the figures mostly from drawings, unless I found the stated explicitly. Comparing P-40 with Fw190 and Yak-1, it looks like that:

 

P-40 - 21.92 -> 22.4

Fw190 - 18.3 -> 17.6

Yak-1 - 17.15 -> 16.2

 

The aircraft lose about 2.3-2.5m² of reference area due to the fuselage, but the elevator sizes are considerably different. Only on the P-40 it is larger than the area lost to the fuselage, so the P-40 does not only lose a smaller percentage of the reference area to the fuselage, but is the only aircraft to have lifting surfaces larger than the reference area.

Bottom line, if the airfoils all came with the same 1.4 lift coefficient over all the surfaces, the coefficients for the reference area would read:

P-40: 1.43

Fw190: 1.35

Yak-1: 1.32

Meaning that just by plane geometry, the lift coefficient of the P-40 would be noticeably larger than those of the Fw190 and Yak-1. For our purposes that is, in game lift calculation is more complex.

 

Given that I'm sceptical towards anything above 1.4, at least for the P-40 I can raise that limit to 1.5 now.

 

I assume that the bold section should read "but the horizontal stabilizer sizes are considerably different" - since you earlier said you will ignore the elevators?

 

This analysis is premised on the hypothesis that the horizontal stabilizers have the same +ve CL as the wings.

 

But do you actually know that the horizontal stabilizers in the P-40 - or the others in the game - produce lift under normal flight or turn conditions?  I thought that the main effect of the tailplane was to balance the GG and centre of lift, in which case they might produce either some lift or downforce, depending on the design and circumstances. But even this is easily achieved with a symmetrical airfoil - it is just about the angle of incidence.  So is an assumption that the horizontal stabilizers, stripped of elevators,  produce the same CLmax as the wing really warranted? 

 

Even if they did, how would this be affected in flight by movements of the elevators? Elevator position could affect any +/- lift from the horizontal stab, in addition to their own pure turning moment.

 

It is an interesting way of looking at the issue but not entirely convincing.

Edited February 20, 2017 by unreasonable

[JtD]

Yes, typo, fixed it. Thanks for that.

 

Tailplane airfoils generally don't have an as high clmax as the main plane, it's just a ballbark I used. The bigger simplification in this regard however is the omission of the elevator. So I wouldn't really worry about the simplification with the horizontal stabilizer.

 

It does have an impact in real life, it's not small and I just wanted to give an idea, not a full calculation, about what it does in this case. If you wanted to do it properly, you would need to investigate the typical elevator and stabilizer angles of attack for the typical aircraft condition near stall and calculate the figures for the full tailplane.

Given the necessary research, it's easily 100 times the effort for a higher accuracy result, still showing the same trend.

[unreasonable]

It is not the detail I am questioning but the basic premise.

 

Obviously it is just a mathematical truism that if you assume tails have any +ve lift then planes with proportionately larger tails will get the larger CLmax calculated using the reference area that does not include it.  But I do not believe that you have demonstrated a trend in actual planes at all - you are just assuming +ve lift from the tail. If the tail was set up to produce a downwards force to balance the CG and CL, a larger tailed plane would end up with a smaller CLmax on the reference area using this arithmetic.

 

In the circumstances of an approach to a Vmin stall or a turn you are certain to have hard back elevator to keep the nose up. Even if the tail as a whole produced lift in level flight, would it in these circumstances - surely what the stabilizer gives as it's AoA increases (by lift, assuming it is not stalled, plus weather vaning) the elevator must be able to take away or you would never be able to turn at all.

 

Obviously a bigger horizontal stab/elevator will exert more turning force, other things being equal. But that is not the issue in the P-40's case AFAIK, no-one has trouble with exerting sufficient turning force, quite the reverse, they stall out easily as the Critical AoA is strangely low: something that is still a bit of a mystery. 

 

(Just got round to asking Han about this and the definition used in the Tech Specs in the Qs for Developers thread: forgot before. Maybe we will get something illuminating). 

[Venturi]

Considering propeller inefficiency or form drag is also very large on the P40 as given by JtD's "efficiency" table - excellent reasoning, that - there most likely is more to the story of the woeful BOM P40 than the damn thing stalling way before it ought to. Although that certainly must be a large component.

[ACG_KaiLae]

I was planning on getting to that, but after #1 We finish going through the clmax stuff and most especially #2, we get the NASM data. Which I'm still waiting on.

[Holtzauge]

JtD: I have to say I’m sceptical about this method of including stabilizer area as a positive contributor to the Clmax like you did in post #688. AFAIK the stabilizer/tail gives a negative contribution to the Clmax in a stable plane. The only time I have seen that the tail gives a positive contribution is if you for example run a wind tunnel test with the tail in a fixed position: In this case the total Clmax of the airplane will be larger but as soon as you trim out the pitching moment in the test, either by changing the elevator angle or the stabilizer angle you end up with a negative contribution. However, if you have some references to support including the stabilizer as a positive contributor like you did above I would be interested to see those.

[JG13_opcode]

Hmm, intuition tells me that the horizontal stab only contributes negative lift while it's at or near the trim point.  Surely at some angle of attack the lift slope crosses into the first quadrant (edit:  unless you think it stalls)?  We should be able to estimate it from lifting-line theory and knowing the free stream velocity.

 

Interesting question

 

edit 2:  spent a few hours googling, can't figure out which airfoil was used on the horizontal stab of the P-40, and my textbooks from university don't have polars for the 2215.  Enthusiasm plummeting, time for a beer 

Edited February 21, 2017 by JG13_opcode

[Holtzauge]

I assume with “moving into the first quadrant” you mean the tail contributing an up force rather than a down force? If so I still remain sceptical since I think you will still have a down force to keep the airplane in trim.

 

Also, I don’t think you can look at the stabilizer in isolation since this is subsonic flow and if you change the angle of the elevator you also change the flow field over the stabilizer. In essence, what you get when you pull back on the elevator is a negatively cambered profile which is combined of the stabilizer and elevator as a unit. In addition, I think the idea that the stabilizer contributes lift since both the wing and the stabilizer have a similar geometric aoa is flawed because the stabilizer/tail sits in the down wash behind the wing so the actual aoa the tail is experiencing is lower than the wing. If you combine this with the negative camber the elevator contributes you will I think end up with a negative contribution to the overall Clmax.

 

At least that is my take on it but as I said before, I may be wrong so I would not mind seeing some references where the tail contributes lift on a stable plane close to the stall but until I do I remain a sceptic.

[JG13_opcode]

I assume with “moving into the first quadrant” you mean the tail contributing an up force rather than a down force?

Sorry, I meant the first quadrant of the polar, as in the region where Cl and alpha are both positive. My thing is that, just as the main wing will generate negative lift at high negative AoA, so too should the horiziontal stab generate positive lift at high positive AoA even if it's negatively cambered. It's basically just an upside down regular airfoil on some aircraft. Again, not knowing what specific section was used for the P-40 I can't say with any greater certainty than that.

Also, I don’t think you can look at the stabilizer in isolation since this is subsonic flow and if you change the angle of the elevator you also change the flow field over the stabilizer. In essence, what you get when you pull back on the elevator is a negatively cambered profile which is combined of the stabilizer and elevator as a unit. In addition, I think the idea that the stabilizer contributes lift since both the wing and the stabilizer have a similar geometric aoa is flawed because the stabilizer/tail sits in the down wash behind the wing so the actual aoa the tail is experiencing is lower than the wing. If you combine this with the negative camber the elevator contributes you will I think end up with a negative contribution to the overall Clmax.

Indeed. In terms of the increased downwash off the wings at high AOA you'll see adverse effects on effective Cl there too. You're probably right. Edited February 21, 2017 by JG13_opcode

[Holtzauge]

Sorry, I meant the first quadrant of the polar, as in the region where Cl and alpha are both positive. My thing is that, just as the main wing will generate negative lift at high negative AoA, so too should the horiziontal stab generate positive lift at high positive AoA even if it's negatively cambered. It's basically just an upside down regular airfoil on some aircraft. Again, not knowing what specific section was used for the P-40 I can't say with any greater certainty than that.

 

Ok, I see what you mean and I think the reason you have to pull up the elevator is just that: As the aoa of the airplane increases, so does the aoa on the tail and if you did not change the "camber" by pulling up elevator the stabilizer would as you say give a positive contribution but in doing so would lift up the tail and then the nose of the plane would go down and you would not reach your stall Clmax. Really, I don't see any way around it: On a stable plane you need a down force on the stabilizer/elevator assembly to lift up the nose.....

 

Actually, looking back a bit I think unreasonable did a good job explaining this "weather waning" effect here in post #695 above.

[JtD]

OK, understood the question now. The necessity to produce lift with the tail at high angles of attack basically comes from two things. One is simple, the other may be not. So I took a bit of time trying to figure out how to best illustrate it.

 

The first thing is the fact that the centre of gravity (CoG) is typically somewhat above the chord of the wing, so pitching up the nose will move the CoG backwards which will necessitate a counter-force from the tail. It might be in the region of 5 cm, but with typical geometry this means that around 1% more of the aircraft weight needs to be supported by the tail near stall than needs to be done at low angles of attack at higher speeds.

 

The second thing is the distribution of the lift along the wing chord, which changes as angle of attack changes. Typically, the WW2 airfoils will have it move somewhat forward up to the stalling point, and significantly back when stalling. A concept to describe these characteristics is the moment coefficient (Cm). Similar to lift coefficient, you multiply it by dynamic pressure, wing area and chord length to arrive at the moment, with which the wing pitches the aircraft. This moment has to be controlled with another moment, such as the forces generated by the tail. The definition is that a positive Cm pitches the nose up.

 

I've attached the characteristics of the three airfoils used on the aircraft I have selected in my above example, and it can be seen on the lower right chart, that all of them have a negative moment coefficient, which means on all of the aircraft the wing wants to pitch the aircraft nose down. Since dynamic pressure is equivalent to the square of speed, this nose down pitching moment is much larger at high speeds and low angle of attack than it is at low speeds and high angles of attack. Or in other words, if at high speeds the plane is balanced, and you go slow towards the stall, the wing nose down pitching moment reduces, the plane goes out of balance, and the nose pitches up. Unless you lift up the tail to counter it.

 

The 230xx series as used on the Fw has a very low Cm, but on the 22xx series used on the P-40 it is considerable. This means the P-40 needs to be trimmed all the time, whereas the Fw190 is relatively stable, and it means that the tail of the P-40 needs to counter far larger forces than that of the Fw190, which in part explains why it is much larger. I don't really know how the Soviets fit that picture, given that the ClarkYH is pretty similar to the 2212 in that regard, while the Yak tail assembly is quite small.

 

I have also attached a calculation by Fw, showing forces on the tail assembly. The first four lines are all medium speed near stall cases, with details depending on centre of gravity position, varying with loadout. You can see that the whole tail has a positive angle of attack, aH is around 16°. The elevator is angled upwards at bH varying between 15..22°. The way I see it bH refers to the chord line of the stabilizer, so with regards to air flow it is near neutral. The net force of the assembly is given with Ph, which is defined as positive if it is a downforce. In the high angle of attack scenarios it's always negative, therefore it is producing between 210kg and 860kg of lift.

 

I hope this explains my point of view sufficiently, if it doesn't, just ask.

 

p.s. I just noticed that all charts also contain the information for the centre of pressure, which is another concept of looking at the aerodynamic balance. To me much easier to understand. It is basically the aerodynamic equivalent to the centre of gravity, and represents all aerodynamic forces in a single point on the wing with a vertical vector. Basically the distance between the upwards CoP force and the downwards CoG force determines the moment which has to be countered by the forces acting on the tail. If you look at the left charts there is a thin C.P. line. You can see that from low angles of attack to higher ones it moves slightly forward along the chord, up to stalling point, where it moves back considerably. Naturally, with the CoP moving forward, the aircraft develops a nose up trend, which will then need to be countered by lift generated by the tail. If there was no lift from the tail, the aircraft would continue raising the nose until it stalls.

Edited February 22, 2017 by JtD

[ZachariasX]

In other words, when the stick is pulled back at high AoA, the elevator is not exerting a downforce, but instead just exerts less lift than it would if the stick is kept neutral.

 

The movement of the pressure point is very evident when you do windsurfing. Older windsurfing rigs (wings, in effect) showed this effect rather drastically, making life hard for the sailor. New rigs are specifically designed to keep the pressure point at the same spot, making it way more controllable and keeping it in trim.

 

I think you raised an important point here JtD, as this also illustrates the need for constant trimming of the P-40 whereas planes with the 230xx profiles do less. Making this impaired control an upside it could help then at high AoA conditions by shifting more lift to the elevator section. Now lts find out how much added lift this could be. So far, this entire trim change is not modelled I guess. I wonder if it was noticeable if it was.

[19//Moach]

In other words, when the stick is pulled back at high AoA, the elevator is not exerting a downforce, but instead just exerts less lift than it would if the stick is kept neutral.

 

 

that REALLY depends on the plane... as far as I recall from my model-glider-building days, the best stability position is not very good for lift/drag - so compromises must be made to attain desired performance without having yourself a deathtrap machine of doom and mayhem

 

curiosity - have you ever noticed how the modern A320 series AirBusses have a wing placed somewhat ahead of the "usual" (noticeable when compared to a 737) - that is actually supposed to make them more fuel efficient, but in doing so, they are somewhat tail-heavy and therefore, unstable - so well, they stick a FBW computer on it and call it a day.... whereas the decades older - and quite more elegant - boeing design actually achieves stability by aerodynamic design (no computers needed, though it gets them anyways by now) ^{[citation needed]}

 

 

anyways - tails that "help" with lift are good for performance - but not so great with stability... and vice versa - there are many workable compromises in between as well as pretty good best-of-both layouts too... I mean, there's planes of all shapes out there, some are even surprising they can be airworthy, with all the weird they got going... but well - when it works it works....

 

 

our poor p-40, however, more doesn't than does "work"... I cannot shake the feeling there is something truly wrong with the dynamics of that plane - something simple, slight, yet brutal in its ultimate effect - minor thought it may be... or maybe a series of small things working together to make for one uncharacteristically scary ride...

 

of all things I've read about the hawks (some less kind than others) - I have not encountered any accounts of it being a dangerous, scary plane to fly - the plane we got, is dangerous - and scary

 

 

if I were a test pilot checking it for a military contract - I'd deem it "unairworthy" at it's current state -- and surely, there were known problems historically (I've read about the whole case) - but if issues had been as severe as we have here, I have serious doubts the plane would have ever been accepted into service with the USAAF (and whatnot)

 

 

I reckon this is a generally "when in doubt, make it bad" style interpretation of the P40's flight dynamics in regards to any available margin for error...

 

 

this plane could definitely use some love, like the 190 just got.... it feels strangely alike, the two... in how they both were, and how the 190 is now, and how the p40 FEELS like it ought to be

 

 

 

but we need more data - so well.... sits and waits

Edited February 22, 2017 by 19//Moach

[Dakpilot]

Regardless of how much I like the P-40, a fighter aircraft that is completely prohibited to enter a spin has some design issues

 

I guess the millions of public money Curtiss used to promote their products during WWII was actually money well spent

 

I also hope that it becomes a bit 'sweeter' to fly in normal flight envelope, with a small tweak to some value like the FW-190, but I also feel many people's expectations of the aircraft are way too high 

 

Cheers Dakpilot

[Cpt_Branko]

Regardless of how much I like the P-40, a fighter aircraft that is completely prohibited to enter a spin has some design issues

 

I guess the millions of public money Curtiss used to promote their products during WWII was actually money well spent

 

I also hope that it becomes a bit 'sweeter' to fly in normal flight envelope, with a small tweak to some value like the FW-190, but I also feel many people's expectations of the aircraft are way too high 

 

Cheers Dakpilot

 

It was not completely prohibited to enter a spin from what I can see.

 

Looking at manuals

P-36 - spins of more than three turns prohibited

P-39 (K/L, Q) - spinning "not recommended"

P-40 (B/C) - no restrictions noted

P-40 (D/E, F-L,  also N) -  spin of more than three turns, spin with baggage, auxillary fuel, or any other overload prohibited

P-51 (Mustang I) - deliberate spinning is not permitted

P-51B/C (Mustang III) - Practice spins are not to be started below 12000 feet. Recovery action is to be taken after no more than two turns.

Spitfire V - Spinning is permitted by pilots who whave written permission from the C.O. of their squadron. Spins are not to be started below 10000 feet, recovery must be started no lower than 5000 feet.

Spitfire XIV - Spinning is permitted. Spins are not to be started below 10000 feet. Recovery is to be initiated before two turns are completed.

P-47 B/C/D - PRACTICE SPINS IN EXCESS OF 1/2 TURN ARE PROHIBITED. (all caps theirs)

F4U - NO INTENTIONAL SPINNING OF MODEL F4U-1, FG-1, F3A-1 AIRPLANES IS PERMITTED (all caps theirs)

F6F - no restrictions noted (earlier RAF manual for Hellcat prohibits spins "pending spinning trials").

Hawker Typhoon - recovery to be initiated after no more than 1 turn 

 

I'm not seeing any particular restrictions on spinning which stand out comparatively speaking when we talk about the P-40. The spin is noted to be violent in one of the manuals, but that's about it. RAF disallowed practice spins on P-40E specifically until 30 gallons of fuel from the fuselage tank have been consumed, and limited practice spins to two turns (which is, if you look above, the same limitation as many other planes).

Edited February 22, 2017 by Cpt_Branko

[=362nd_FS=Hiromachi]

 

 

I'm not seeing any restrictions which are out of the ordinary.

Hey Branko, nice to see you around !

It's pretty standard thing that warns about or even prohibits spinning. Back then and today. I could find the same warning in Ki-43/Ki-84 manual from 1943-1944 and in the recommendations for Flying The BF 109 G by Dave Southwood from November 2002 for their 109 G "Black 6". 

[Cpt_Branko]

Likewise

 

Anyway, yeah, it seems pretty common. The few planes which don't limit practice spinning to some extent (normally 1-2 turns) actually "stand out".

 

Really, we need someone to volounteer to spend a 100-150$ on getting documents from the British, and our understanding of the capabilities of the aircraft would be much improved.

Edited February 22, 2017 by Cpt_Branko

[Dakpilot]

I am always happy to be proven wrong

 

there have been many sections of official manuals posted already stating that spins are prohibited in P-40 models

 

but there are many notes that the P-40's had issues from needing the longer tailwheel to sort out  landing issue, trimming problems when diving, the Russian "tumble" and the fuselage extensions 

 

Compared to many designs the P-40 series seems to have had needed a lot of adjustment, not that these are excuses for current FM issue but I personally feel it was a flawed design when compared to many others,

 

see this post

 

https://forum.il2sturmovik.com/topic/25323-p-40-turn-rateflight-model-check/?p=425049

 

anyway hope further info is found to improve P-40E

 

Cheers Dakpilot

[ZachariasX]

that REALLY depends on the plane... as far as I recall from my model-glider-building days, the best stability position is not very good for lift/drag - so compromises must be made to attain desired performance without having yourself a deathtrap machine of doom and mayhem

 

[...] snipp

 

but we need more data - so well.... sits and waits

It is my understanding that all planes that we have in-game have a performance envelope defined by thrust, weight, and the lift-drag polar. In this sense, I see absolutely nothing other than the mathematical result of the best effort the dev team can do to get these parameters right and compute the performance using lifting line theory, but no intentional bias whatsoever.

 

With the P-40, it seems to me that we are elucidating an effect happening near the edge of the performance enveloppe at high AoA. As max. substained turn is restricted by thrust, the effect of the shifting pressure point of the wing should not be noticeable at all. Thus by introducing and simulating this effect, the P-40 is not expected to creep up in a circle on a Bf-109 at that speed if one just parks it there "and waits for the target to appear in the gunsight."

 

But it should help in instataneous maneuvrability meaning one could get just a bit more alpha for a snapshot on a target.

 

What I find interessting is that should be a so far missing effect on all planes that one might thinkf adding, as it surely would add or discount some aspects of the combat performance. It would also make the requirements for constant re-trim apparent. Especially this later effect would influence the ease with which a plane can be flown in its optimal state. Suddenly people would notice the difference of a "benign" aircraft and a quirky one much more. So far, this difference is mainly about how sharp the stall is.

 

And yes, Americans were and are great salesmen. But so was Willy Messerschmitt.

[ZachariasX]

 

Here is an illustration of the movement of the lift vector at increasing AoA on a ClarkY profile showing a significant forward move. (I hope it is a scientifically sound diagram, but it looks right at the first glance.)

 

EDIT: The problems of a shifting centre of pressure on a wing at varied AoA is in similar fashion discussed in an exposé from StanfordU http://adg.stanford.edu/aa241/stability/staticstability.html or NASA. https://www.grc.nasa.gov/WWW/K-12/airplane/cp.html

Edited February 22, 2017 by ZachariasX

[Holtzauge]

I had a look at the Fw table JtD posted and to me it looks like it calculates tail loads associated with different load cases to be used for dimensioning purposes. The dynamic pressure used is also much higher than at stall speeds q=709 kg/m**2 corresponds to roughly 383 Km/h IAS and q=3500 kg/m**2 to 850 Km/h so way above stall speeds and also it looks like there is a substantial load factor (ntr?) in many cases . So while the tail load may be 210 to 860 Kg lift in the first 4 lines in the table this is not at the 1 g stall IAS at around 180 Km/h but at substantially higher IAS and at what looks to me like a load factor of between 4.1 to 6 but since the figure is a bit fuzzy I’m assuming ntr is load factor. So as far as I can tell the table tells us little about the tail load at stall.

 

Anyway, another way of looking at this is to do a ballpark sanity check of a probable tail load: Assuming a weight of 3850 Kg for a Fw-190, a moment arm of around 5.5 m for the tail to the c.g. and eyeballing a moment arm of around 0.1 m for the wing (allowable c.g. range assumed to be about 0.5 to 0.7 m from wing root leading edge). Also note that this is not even including the Cm from the profiles JtD posted which is negative and would need a downforce on the tail to be balanced and in addition, I’m even assuming that the wing lift is so far forward of the c.g. which I don’t think is the case IRL. Using these numbers you end up with a lifting force of around 70 Kg to balance the wing which is a lift contribution of under 2%.

 

Now instead try entering a lifting force of around 200 to 800 kg from the tail and then calculate where the wing needs to be placed in relation to the c.g to balance the resulting 3650 or 3050 Kg respectively the wing produces assuming such a tail load: In the case of the 800 kg stabilizer lift the wing needs to be 1.4 m ahead of the c.g……..

 

In addition and as I said before: I think looking at the data for wing profiles in isolation does not give the whole picture: You also need to take the downwash effect on the tail from the main wing into account and even if this in combination would result in a bit of lifting force on the tail as far as I can see it will be close to negligible or else you will get the weather wane effect unreasonable mentioned earlier.

 

So in summary, to me it seems misleading to include the stabilizer area to make a new reference area which is bigger than the original P-40 wing reference area since even if the tail produces lift it would be a very small contribution (even if the stabilizer has large dimensions) or the plane would simply weather wane.

 

Caveat: It’s been a busy evening for me so I have not had the time to think this thorough properly or recheck calculations but these are my ideas at the time of writing anyway…

 

[JtD]

Nothing wrong with your calculation, if you divide the forces by the load factor you end up with tail lift contribution of 1-4% with the Fw table. The main point was to show that the tail near stall produces a net uplift, not a downforce and to explain the mechanics behind that. We could try to estimate some numbers, but having tried that already, it's very difficult to be sufficiently accurate, given the lack of detailed input data.

 

Please keep in mind that I excluded the elevator effects from my estimate, while the Fw numbers include it. The uplift from the stabilizer alone is significantly larger than what is shown in the table, plus the main point was to show that reference area is not directly comparable between different aircraft. With the P-40 getting a bonus out of that, and that's true even if you just exclude the fuselage area.

 

My personal conclusion was that if the Fw190/Yak-1 are in the 1.35 region, the P-40 would be in the 1.40-1.45 region. According to the charts I posted, the 2212 actually achieves a higher clmax @ ~RN3000000 than both the ClarkYH and the 23012, making figures as high as 1.45 somewhat plausible.

[Holtzauge]

OK, then we are pretty much on the same page I think: Since the c.g. and wing MAC are probably quite close together whatever the contribution from the tail it will most likely at the stall be quite small or else the forces around the c.g. don't balance.

 

Some additional ideas: I'm not sure the Fw-190 table can be used to deduce actual tail loading as in a stable flight condition since if the table is to be used for tail dimensioning as in structural engineering the values in the table could well be those that the tail experiences when you initiate the load factors in the table: For example, with a given weight and external stores loadout the pilot initiates a 6 g manouver and in order to do that he deflects the elevator causing the aoa and elevator angle in the table which at the given dynamic pressure q gives a transient load like listed in the table. Not saying that is what the table is showing since I don't have the other pages but for doing the structural engineering on the tail you need that kind of data.

 

As to the P-40 having a Clmax of 1.4-1.45 (based on a standard reference area), I still think that sounds optimistic since the Me-109 which lacks wing twist (and can thereby load the outer parts more when the root stalls than a wing with washout) only manages 1.4.

 

Anyway, hopefully we will get some measured Clmax data from the archives but until then I’m still leaning towards a Clmax for the P-40 in the same range as most other WW2 fighters, i.e. in the range of 1.3 to 1.4.

 

@ZachariasX: Thanks for the Stanford link about stability and the contribution of different parts of the airplane: Always nice to read something about a complex subject that is so well explained and accessible!

[ZachariasX]

Just trying to ballpark the force we are potentially dealing with.

 

If I make up an aircraft that weights 30 kN and that has his weight distributed 20 kN at a distance of 2 m in front of the Cg (centre of gravity) as well as 10 kN 4 m behind the Cg. For simplicity I assume that all weight is at the centre of gravity of the front and the tail arm respectively, so that shifting the Cg does not affect weight distribution.

 

With that, front part and tail section both are in balance, exerting a torque of (20 kN * 2 m, 10 kN * 4 m) 40 kN/m and thus are in balance. Looking at the diagram depicting the shift of the pressure point on a ClarkY, I assume a 0.5 m shift forward of the original Cg at max. AoA.

 

Shifting the balance like that results the engine arm to exert a torque of only (20 kN * 1.5 m) 30 kN/m. The tail arm however exerts (4.5 m * 10 kN) 45 kN/m.

 

To put it back in balance, the tail has to compensate a torque of 15 kN/m. If I place the elevator section 6 m behind the (shifted) Cg, then it has to LIFT 2.5 kN.

 

In conclusion, my aircraft that is in balance at max. AoA has 8.3% of it weight distributed on the horizontal stabilizer (regardless of its size).

 

Based on this, I can infer that the "stubbier" the aircraft is (like an I-16) the more this effect comes into play, as the shift of the planes Cg is larger in relation to the distance of the Cg of the front part to the Cg of the tail section of the plane. This way it becomes VERY apparent that the Fw-190 D became much a more stable compared to the A variants, greatly mending its stability issues at high altitude. This at the price of a decreased maneuvability. The Dora turned less good than the Anton (instantaneous), despite weight and power advantage.

 

Taken together, as for now I would consider an increase in lift up to maybe an added 10% plausible.

[Holtzauge]

Sure, using the assumption that the c.g. moves as far as you do then maybe 10% is plausible. However, looking at the Yak-1 which has the Clark YH which is similar to the Clark Y figure you posted here, the Yak-1 has AFAIK a typical c.g. range in the order of 23 to 28% MAC and if you trace that into the figure, you can see that that roughly corresponds to a position slightly ahead of the purple vector for 13.4 deg aoa to roughly the yellow 10.8 deg aoa vector. So basically the c.g. and wing lift vectors coincide and in this case the ballpark estimate will instead be that the tail load will be close negligible.

[ZachariasX]

Sure, using the assumption that the c.g. moves as far as you do then maybe 10% is plausible. However, looking at the Yak-1 which has the Clark YH which is similar to the Clark Y figure you posted here, the Yak-1 has AFAIK a typical c.g. range in the order of 23 to 28% MAC and if you trace that into the figure, you can see that that roughly corresponds to a position slightly ahead of the purple vector for 13.4 deg aoa to roughly the yellow 10.8 deg aoa vector. So basically the c.g. and wing lift vectors coincide and in this case the ballpark estimate will instead be that the tail load will be close negligible.

Such a position is a good choice to keep the plane in an aerodynamically stable. Even for R/C models I (over the thumb) check Cg in roughly the 25% area of the wing to start out with. Once one knows the plane you can get incrementally more tail heavy, depending on what you want to do.

 

One question: Does BoX model power on CLmax as well? Or is it always power off? Could it be that in a power on state propwash would create added lift from the rather large elevator of the P-40?

[unreasonable]

Just a couple of thoughts while trying to digest this most interesting discussion:

 

 1) If the wing is at high AoA, then so is the fuselage. This will also provide a rotational force (not lift) - how much I suppose depending on the distribution of the fuselage area compared to the CG. So, would this net out as a force tending to raise the nose, or to raise the tail? I would have thought usually the latter.  Is it a material effect? 

 

2) We do not know how much of this (all the stuff discussed above) is modeled in the FM but I would have thought that much of it must be. Hard to imagine that the FM has the CG and CL at the same point for instance: if they are separate surely the effects of shift in CG with AoA changes must be included automatically?

 

But it is odd to speculate about corrections to a model of which we know not the details. Should someone not just ask Han if the FM models lift at the tail surfaces explicitly? If it does, is this discussion not moot?

[Holtzauge]

Such a position is a good choice to keep the plane in an aerodynamically stable. Even for R/C models I (over the thumb) check Cg in roughly the 25% area of the wing to start out with. Once one knows the plane you can get incrementally more tail heavy, depending on what you want to do.

 

One question: Does BoX model power on CLmax as well? Or is it always power off? Could it be that in a power on state propwash would create added lift from the rather large elevator of the P-40?

 

Yes, I remember that rule from my model flying days as well: Only no RC for me: Had to make do with 2.5 cc diesel powered line controlled combat models.....

 

About the effect of power on Clmax: I don't know but I would  assume that it comes with the package? I mean AFAIK they do model the slipstream effects and the lift vector gets tilted as well with the aoa.

 

Just a couple of thoughts while trying to digest this most interesting discussion:

 

 1) If the wing is at high AoA, then so is the fuselage. This will also provide a rotational force (not lift) - how much I suppose depending on the distribution of the fuselage area compared to the CG. So, would this net out as a force tending to raise the nose, or to raise the tail? I would have thought usually the latter.  Is it a material effect? 

 

2) We do not know how much of this (all the stuff discussed above) is modeled in the FM but I would have thought that much of it must be. Hard to imagine that the FM has the CG and CL at the same point for instance: if they are separate surely the effects of shift in CG with AoA changes must be included automatically?

 

But it is odd to speculate about corrections to a model of which we know not the details. Should someone not just ask Han if the FM models lift at the tail surfaces explicitly? If it does, is this discussion not moot?

 

Good point about the fuselage: As you say it should provide a rightening moment but not much in terms of lift other than the part close to the wing of course. About the c.g. and Cl: I would assume that is rather well modeled with a distribution of weight to model inertia and vertical c.g. location etc or otherwize it would be difficult to get any semblance of realistic ground handling and if some sort of panel model is used for lift this should capture Cl effects as well. OTOH when you get close to the stall and start getting flow separation then I would assume that the panel method does not do very well and they probably have some sort of tweaking of the model here but of course that is pure speculation.

 

About if the tail lifts or not I think the panel model would at least capture parts of this because effects like downwash on the tail would come with the package but when it comes to other effects like the way the lift vector moves on the wing in the figure ZachariasX posted above I'm not so sure.

 

Anyway, what bought on this whole discussion was if the tail on the P-40 in being bigger would produce a significant contribution in lift thus complementing the wing resulting in a higher effective Clmax. So far I have to say that to me that does not seem to be the case since the wing lift vector at stall is so close to the c.g that any significant tail lift would simply weather wane the plane just like you pointed out earlier. As far as I can see the only effect at Clmax the larger stabilizer on the P-40 will have is that it will provide a lower tail Cl loading since it has a larger surface area. IMHO I find it more likely that the reason the P-40 has such a large stabilizer is probably more in terms of providing enough control authority in some load out scenario with the c.g. shifted due to external stores rather than to provide a boost in lift.