Dora vs Mustang: Turning

[Hummingbird]

You do not take in account different CL especially at the speed the best turn ratio is. Mustang has better CL max at low M-numbers and much preferrable at medium.

 

The Mustang doesn't have a better CL max at any speed, the laminar flow airfoil simply won't allow it. The FW190's NACA 23xxx airfoil features a higher CL max and critical AoA thanks to its wider leading edge - whilst the Mustangs narrow LE reduces the CL max & critical AoA, esp. at lower speeds.

[Barfly]

But the Mustang wing has a better drag profile at the speed ranges we're talking about, and that factors into best sustained turn rate.

[Kurfürst]

Lots of speculations on Cl and wing profile characteristics, but what is needed is hard aerodynamics data...

[Yo-Yo]

The Mustang doesn't have a better CL max at any speed, the laminar flow airfoil simply won't allow it. The FW190's NACA 23xxx airfoil features a higher CL max and critical AoA thanks to its wider leading edge - whilst the Mustangs narrow LE reduces the CL max & critical AoA, esp. at lower speeds.

 

Please, prove your statement with something more than a bare words. I know this tale about Mustang airfoil but the test reports tell very different.

[Hummingbird]

Please, prove your statement with something more than a bare words. I know this tale about Mustang airfoil but the test reports tell very different.

 

That's easy Yo-Yo, take a look at the stalling speeds of the aircraft in question, that will tell you all you need to know about the CL max of either aircraft. It's easily calculated. The similar stalling speeds of both aircraft under the same conditions prove beyond any doubt that the 190's wing provided a higher CL max, it quite simply had to because of the higher wing loading.

 

The laminar flow airfoil was designed for optimum performance at low AoA's, featuring a noticable decrease in profile drag over conventional airfoils, it hower payed a price at higher AoA's where boundary layer seperation happened sooner and more suddenly.

[Narushima]

P-51D airfoil:

Tip: http://airfoiltools.com/airfoil/details?airfoil=p51dtip-il

Root: http://airfoiltools.com/airfoil/details?airfoil=p51droot-il

 

FW 190D airfoil: Couldn't find it. The closest I could find was a similar wing with thinner chord

Root: http://airfoiltools.com/airfoil/details?airfoil=naca23015-il

[Yo-Yo]

That's easy Yo-Yo, take a look at the stalling speeds of the aircraft in question, that will tell you all you need to know about the CL max of either aircraft. It's easily calculated. The similar stalling speeds of both aircraft under the same conditions prove beyond any doubt that the 190's wing provided a higher CL max, it quite simply had to because of the higher wing loading.

 

The laminar flow airfoil was designed for optimum performance at low AoA's, featuring a noticable decrease in profile drag over conventional airfoils, it hower payed a price at higher AoA's where boundary layer seperation happened sooner and more suddenly.

 

Complicated questions always have simple, obvious but wrong answers... :), try again, please. It's not a proof, only common words again.

Edited August 28, 2014 by Yo-Yo

[Pilum]

In NACA report 829 which investigates Cl max for different a/c configurations, “airplane 1” is the P51 and it actually has a Clmax of 1.4 according to this report. Note that this is the Clmax for the wing and that while Clmax figures for the wing profile as such may be interesting for comparing different profiles, they cannot be assumed to be automatically representative of the whole wing.

 

In this context the 1.4 figure for the P51 is surprisingly high, since the Me109 according to full scale tests has 1.4-1.45 while the Spitfire has around 1.36.

 

It should be noted that this is all low Mach Clmax and that Clmax is Mach/speed dependant. While the effect is naturally greater at higher alt (due to higher turn speeds) and at accelerated turns, it is also present at best sustained turn rates at low altitude.

 

I checked my C++ assumptions and I have the Clmax for the Dora at 1.35 at M=0.2, 1.31 at M=0.3 and 0.8 at M=0.6. Can’t remember where I got the 1.35 from, but it shown the trend anyway.

 

If anyone has some better data and references on the P51 and Dora, that would be interesting. A Clmax of 1.4 for the P51 seems a bit on the high side.

 

Yo-Yo: Could you give us some input on what Clmax you have assumed for the P51 and Dora and how Mach influences this in your model?

[wolfstriked]

In the other sim that we are not allowed to compare DCS to the plane side slips but as speed increase the ball stays off to the side.I asked on their forums and was told by a dev that this is how real planes work and speed does not influence the ball position at all.This seemed so silly to me as common sense is that the vertical stabilizer is a wing and that even with no rudder input it will want to straighten out as speed increases.

 

In IL2-46 I remember a build that had it so that as speed increased the ball would slowly center and then stay there until speed decreased.Now in DCS WW1 planes the ball does move with speed but it keeps going past center and into the opposite direction with speed.Is that something that can be fixed or is that the way its supposed to work.As I said many times I am probably wrong and DCS is right.;)

[Hummingbird]

Pilum,

 

The FW190's CL max was 1.58, taken directly from Focke Wulf documents.

 

The 109's without slats was 1.48 (as you said), and 1.70 with slats deployed.

 

As for the P-51, I have 1.38-1.40 from a NACA doc as well, but this assumes a completely smooth wing (important for the laminar flow type airfoil), whilst the German figures were extrapolated from operational aircraft.

Edited August 28, 2014 by Hummingbird

[Hummingbird]

Complicated questions always have simple, obvious but wrong answers... :), try again, please. It's not a proof, only common words again.

 

There is no better proof Yo-Yo.

 

But if you're only interested in the theoretical then so be it, the 190's CL max is readily available on FW AG documents, whilst the P-51's is at hand in NACA documents.

 

Finally let me quote a previous post here:

"1940's windtunnel tests were nortoriously inaccurate"

 

Hence I tend to rely mostly on real world collected data such as stalling speeds & take off distances rather than some theoretical windtunnel figures ;)

Edited August 28, 2014 by Hummingbird

[Pilum]

Pilum,

 

The FW190's CL max was 1.58, taken directly from Focke Wulf documents.

 

The 109's without slats was 1.48 (as you said), and 1.70 with slats deployed.

 

 

Can you source the 1.58 for the Fw190? That sounds high.

 

The only 1.7 Clmax figure I have seen is in a Me109F aerodynamics summary doc that refers to landing configuration so with flaps deployed. Can you source the 1.7 figure please?

[gavagai]

Speaking of more words, here is one to try out:

 

A purely laminar flow wing, i.e. no turbulent airflow, would stall with any non-zero angle of attack. That one I heard from a friend who was taking a physical chemistry course with the USAF.

 

Here is a good resource on airfoils, and they have data for both the wing root and wingtip of the P-51.

[OutOnTheOP]

In the other sim that we are not allowed to compare DCS to the plane side slips but as speed increase the ball stays off to the side.I asked on their forums and was told by a dev that this is how real planes work and speed does not influence the ball position at all.This seemed so silly to me as common sense is that the vertical stabilizer is a wing and that even with no rudder input it will want to straighten out as speed increases.

 

In IL2-46 I remember a build that had it so that as speed increased the ball would slowly center and then stay there until speed decreased.Now in DCS WW1 planes the ball does move with speed but it keeps going past center and into the opposite direction with speed.Is that something that can be fixed or is that the way its supposed to work.As I said many times I am probably wrong and DCS is right.;)

 

Fairly certain that's correct behavior: the vertical stabilizer for some aircraft of the period (I'm fairly certain including the Mustang) was twisted slightly to one side, so there was a "zero trim" speed that it was optimized for. At that speed, the aerodynamic forces acting on the tail would be perfect to balance out the engine torque and propwash effects- I would imagine most aircraft had this set for cruise speed (and on the Mustang, it seems to be around 250-270 mph, right around cruise speed).

 

Below that speed, the engine torque (and effects of propwash) push the tail one direction, while above that speed, the slipstream pressure on the vertical stabilizer overcomes (and then exceeds) those forces and pushes the tail the other direction.

 

There's a reason they had rudder trim!

[Yo-Yo]

There is no better proof Yo-Yo.

 

But if you're only interested in the theoretical then so be it, the 190's CL max is readily available on FW AG documents, whilst the P-51's is at hand in NACA documents.

 

Finally let me quote a previous post here:

"1940's windtunnel tests were nortoriously inaccurate"

 

Hence I tend to rely mostly on real world collected data such as stalling speeds & take off distances rather than some theoretical windtunnel figures ;)

 

Let's play a court (non-martial, for sure!) - you stated something and you have to prove this statement, not me.

 

AS you find something it will be interesting, but be careful - I have at least one ace in my sleeve...

[wolfstriked]

Fairly certain that's correct behavior: the vertical stabilizer for some aircraft of the period (I'm fairly certain including the Mustang) was twisted slightly to one side, so there was a "zero trim" speed that it was optimized for. At that speed, the aerodynamic forces acting on the tail would be perfect to balance out the engine torque and propwash effects- I would imagine most aircraft had this set for cruise speed (and on the Mustang, it seems to be around 250-270 mph, right around cruise speed).

 

Below that speed, the engine torque (and effects of propwash) push the tail one direction, while above that speed, the slipstream pressure on the vertical stabilizer overcomes (and then exceeds) those forces and pushes the tail the other direction.

 

There's a reason they had rudder trim!

Hmmmm interesting and that makes sense,thanks.:D

[Hummingbird]

Can you source the 1.58 for the Fw190? That sounds high.

 

The only 1.7 Clmax figure I have seen is in a Me109F aerodynamics summary doc that refers to landing configuration so with flaps deployed. Can you source the 1.7 figure please?

 

Yes.

 

For the FW data:

 

 

 

 

As for the 109, you're right there is a document which lists 1.7 with flaps AND wheels down IIRC, but wheels down would also significantly decrease the CL max.

 

CL max wheels and flaps up, no slats, for the 109 should be 1.48, right? Now in the areas covered by the non camber increasing slats the increase to the lift coefficient should be by a magnitude of approx. 35% - if we are to go by this NACA chart for the Clark Y airfoil (IIRC the 109's 2R1 airfoil was a modified Clark Y airfoil with a small increase in lift):

 

 

So an overall CL max of 1.65-1.7 seems reasonable for the 109.

[Hummingbird]

Let's play a court (non-martial, for sure!) - you stated something and you have to prove this statement, not me.

 

AS you find something it will be interesting, but be careful - I have at least one ace in my sleeve...

 

Please do post your ace :)

[Yo-Yo]

Please do post your ace :)

 

Sorry, what CL do you consider as a CLmax for FW190 clean wing?

 

Do you think that this number is right for the speed the plane turns at?

[Pilum]

Yes.

 

For the FW data:

 

 

 

 

As for the 109, you're right there is a document which lists 1.7 with flaps AND wheels down IIRC, but wheels down would also significantly decrease the CL max.

 

CL max wheels and flaps up, no slats, for the 109 should be 1.48, right? Now in the areas covered by the non camber increasing slats the increase to the lift coefficient should be by a magnitude of approx. 35% - if we are to go by this NACA chart for the Clark Y airfoil (IIRC the 109's 2R1 airfoil was a modified Clark Y airfoil with a small increase in lift):

 

 

So an overall CL max of 1.65-1.7 seems reasonable for the 109.

 

First of all let me say that you can seldom from wing profile Clmax data (2D data) like you have posted make accurate predictions how much the Clmax for a whole wing will be because there are elements of span loading and boundary layer transport in the spanwize direction that make the 3D analysis and behaviour of the wing as whole difficult to predict. In addition, most wings have different profile thickness, twist and even profiles at the root and at the tip.

 

Beginning with the data you have posted on the Fw190, the 1.58 Clmax figure you have quoted is from the section detailing the take off and landing aerodynamics: Note that the take off condition Clmax is as you quite correctly quouted 1.58 but that this is when the flaps are deployed 12 degrees. So if this was your proof of the clean wing Clmax=1.58 then I'm afraid you are mistaken.

 

When it comes to the Me109 figure of 1.7 it is now clear where that comes from. You have mistakenly assumed that the wing profile Clmax can be applied to the whole wing: The problem here is that while the parts of the wing with the slats can retain attached flow to a much higher angle of attack (aoa) than the rest of the wing this can never be utilized because the rest of the wing will stall at a significantly lower aoa. The reason you put part span slats on a plane like on the 109 is to retain roll control at higher aoa so you avoid a departure in stalling conditions. The slats of the Me109 start to deploy at around 10 deg aoa and are fully deployed at around 20 degrees. However, the full Clmax=1.7 on the slatted part of the wing is not attained until around 30 degrees by which time the rest of the wing is hopelessly stalled.

 

So, the object of the slats is simply to ensure that roll control is maintained. The maximum Clmax of the wing as a whole is reached at around 20 degrees when the unslatted part of the wing stalls. The Clmax of the Me109 has been measured both by British RAE and by the Germans themselves in the Charles Meudon wind tunnel. They both agree the Clmax for the wing as such (not wing section data like you posted!) is around 1.4. This is of course for the Me109E but if you for some reason think the Me109G and K will be significantly different in this respect and have a clean wing Clmax of 1.7 then please post some valid proof of this because you cannot expect to be taken seriously if you rest your case on the 2D wing profile data and an aoa of 30 degrees as you have done so far.

[Hummingbird]

Alright Pilum, since you want to go into details.

 

 

Beginning with the data you have posted on the Fw190, the 1.58 Clmax figure you have quoted is from the section detailing the take off and landing aerodynamics: Note that the take off condition Clmax is as you quite correctly quouted 1.58 but that this is when the flaps are deployed 12 degrees. So if this was your proof of the clean wing Clmax=1.58 then I'm afraid you are mistaken.

 

Where does it say that the 1.58 figure on this document is for landing configuration? I see this nowhere.

 

When it comes to the Me109 figure of 1.7 it is now clear where that comes from. You have mistakenly assumed that the wing profile Clmax can be applied to the whole wing: The problem here is that while the parts of the wing with the slats can retain attached flow to a much higher angle of attack (aoa) than the rest of the wing this can never be utilized because the rest of the wing will stall at a significantly lower aoa. The reason you put part span slats on a plane like on the 109 is to retain roll control at higher aoa so you avoid a departure in stalling conditions. The slats of the Me109 start to deploy at around 10 deg aoa and are fully deployed at around 20 degrees. However, the full Clmax=1.7 on the slatted part of the wing is not attained until around 30 degrees by which time the rest of the wing is hopelessly stalled.

 

So, the object of the slats is simply to ensure that roll control is maintained. The maximum Clmax of the wing as a whole is reached at around 20 degrees when the unslatted part of the wing stalls. The Clmax of the Me109 has been measured both by British RAE and by the Germans themselves in the Charles Meudon wind tunnel. They both agree the Clmax for the wing as such (not wing section data like you posted!) is around 1.4. This is of course for the Me109E but if you for some reason think the Me109G and K will be significantly different in this respect and have a clean wing Clmax of 1.7 then please post some valid proof of this because you cannot expect to be taken seriously if you rest your case on the 2D wing profile data and an aoa of 30 degrees as you have done so far.

 

First of all the Charles Meudon wind tunnel test was with a wing without slats and a reduced wing aspect ratio, both of which are factors that influence the CL max.

 

Secondly the measured CL max at Charles Meudon was 1.48. With slats this figure will increase, as the thicker inboard part of the wing features a higher CL max than the thinner outboard part. Thus expecting an overall CL max of 1.4 is to be even more mistaken.

 

By comparison the P-51B's CL max is listed as 1.28 in NACA report 829, page 26. I wouldn't expect the P-51D's to be any higher.

 

Finally the objective of the slats was to increase the overall lift provided by the wing, thereby improving maneuverability and lowering the landing speeds. Anything else wouldn't have justified the extra complexity and a simple washout would've been chosen.

 

Again from NACA report 829, you can here see the effect of outboard mounted slats on a wing planform very similar to the 109's (infact I believe this was meant to be a copy):

 

Note the CL max of ~1.8.

Edited August 31, 2014 by Hummingbird

[Yo-Yo]

Alright Pilum, since you want to go into details.

 

 

 

Where does it say that the 1.58 figure on this document is for landing configuration? I see this nowhere.

 

 

 

First of all the Charles Meudon wind tunnel test was with a wing without slats and a reduced wing aspect ratio, both of which are factors that influence the CL max.

 

Secondly the measured CL max at Charles Meudon was 1.48. With slats this figure will increase, as the thicker inboard part of the wing features a higher CL max than the thinner outboard part. Thus expecting an overall CL max of 1.4 is to be even more mistaken.

 

By comparison the P-51B's CL max is listed as 1.28 in NACA report 829, page 26. I wouldn't expect the P-51D's to be any higher.

 

Finally the objective of the slats was to increase the overall lift provided by the wing, thereby improving maneuverability and lowering the landing speeds. Anything else wouldn't have justified the extra complexity and a simple washout would've been chosen.

 

Again from NACA report 829, you can here see the effect of outboard mounted slats on a wing planform very similar to the 109's (infact I believe this was meant to be a copy):

 

Note the CL max of ~1.8.

 

THe curves you posted have CL at AoA 0 about 0.6. It's absolutely impossible for the clean wings because, for example, that the plane must fly nose down attitude even at climb speed.

 

Typical CL_0 for the fighter plane is about 0.1 in clean configuration.

[Pilum]

Alright Pilum, since you want to go into details.

 

 

 

Where does it say that the 1.58 figure on this document is for landing configuration? I see this nowhere.

 

 

 

First of all the Charles Meudon wind tunnel test was with a wing without slats and a reduced wing aspect ratio, both of which are factors that influence the CL max.

 

Secondly the measured CL max at Charles Meudon was 1.48. With slats this figure will increase, as the thicker inboard part of the wing features a higher CL max than the thinner outboard part. Thus expecting an overall CL max of 1.4 is to be even more mistaken.

 

By comparison the P-51B's CL max is listed as 1.28 in NACA report 829, page 26. I wouldn't expect the P-51D's to be any higher.

 

Finally the objective of the slats was to increase the overall lift provided by the wing, thereby improving maneuverability and lowering the landing speeds. Anything else wouldn't have justified the extra complexity and a simple washout would've been chosen.

 

Again from NACA report 829, you can here see the effect of outboard mounted slats on a wing planform very similar to the 109's (infact I believe this was meant to be a copy):

 

Note the CL max of ~1.8.

 

First of all, concerning the Fw190, I said start configuration not landing. The flap angles are listed 7-8 rows lower than where you read off the Camax value. Another plausible interpretation of the whole table section is that it all relates to the propeller Cl and Cd (Ca and Cw in German nomenclature). Anyway, the 1.58 just hangs in there and I cannot see how you can cherry pick it out of that table and claim it is the clean wing Clmax. Could be wing Clmax with flap angle, could be propeller Clmax. Take your pick. If this is your basis to claim 1.58 clean wing Clmax for the Fw190 you don’t convince me and I’m sure you won’t convince the developers either.

 

Concerning the figure 30 from NACA report 829 you refer to as a setup you believe to be “meant as a copy of the 109” you are similarly mistaken: It is for airplane 7 which is depicted on page 2 and AFAIK is a Curtiss A-25 Shrike dive bomber. As Yo-Yo pointed out, you can immediately tell that this chart is not for a fighter configuration due to the high Cl at zero aoa. How you can refer to this as something you find representative for the Me109 is beyond me. In addition, the text specifically mentions that the Clmax was not improved by the slats and shows that the improved slat maintained the flow over the aileron part of the wing so that roll control is maintained just like i said in my earlier post. Anyway, this data is moot in the discussion since it presents data on the Curtiss A-25 Shrike bomber, not the Me109 fighter.

 

Getting back to your theories on improvements due to slats and how these can improve Clmax, let me repeat the explanation I posted before in more explicit terms: A conventional clean wing will stall at around 20 degrees giving a Clmax of 1.2-1.5. Adding a slat means that the wing profile behind the slat has the potential go beyond 20 degrees aoa up to about 30 degrees and at this aoa attain maybe 1.7 to 1.8. So, if you add a slat that spans the whole wing you can theoretically reach 1.7-1.8 at 30 degrees aoa. However, a part span slat will only help the part of the wing behind the slat. The rest will stall and loose lift at around 20 degrees aoa meaning the max Cl you can take out of the wing occurs at about 20 degrees, i.e around 1.2-1.5. So apparently it bears repeating: the slats on the Me109 are there to retain roll control at stall, not to improve Clmax. In fact, the Me109 wing does not incorporate a twist (wash out) like many other designs to retain roll control making the slats all the more important in this respect. I hope this makes it more clear how slats work and why they were included on the Me109.

 

So if you want to continue to push for a clean wing Clmax of 1.58 for the Fw190 and 1.7 for the Me109 in DCS it looks like you still have some ways to go.:smilewink:

Edited August 31, 2014 by Pilum

[Hummingbird]

First of all, concerning the Fw190, I said start configuration not landing. The flap angles are listed 7-8 rows lower than where you read off the Camax value. Another plausible interpretation of the whole table section is that it all relates to the propeller Cl and Cd (Ca and Cw in German nomenclature).

 

Not quite as, first of all then you have two different CL max (Ca max) figures for the same prop, which can't be the case.

 

Secondly the flap angle figures simply show the angles used for landing & start, nowhere does it say or indicate that the Ca max is for any of these settings.

 

Furthermore the F4U Corsair which sported the same NACA 23XXX airfoil featured an identical CL max of ~1.6 according to NACA report 829. Thus according to all available evidence the 1.58 figure is for clean wing configuration

 

 

Anyway, the 1.58 just hangs in there and I cannot see how you can cherry pick it out of that table and claim it is the clean wing Clmax. Could be wing Clmax with flap angle, could be propeller Clmax. Take your pick. If this is your basis to claim 1.58 clean wing Clmax for the Fw190 you don’t convince me and I’m sure you won’t convince the developers either.

 

Again, read NACA report 829, the Corsair with the same airfoil features a near identical clean wing CL max. That is proof enough.

 

Getting back to your theories on improvements due to slats and how these can improve Clmax, let me repeat the explanation I posted before in more explicit terms: A conventional clean wing will stall at around 20 degrees giving a Clmax of 1.2-1.5. Adding a slat means that the wing profile behind the slat has the potential go beyond 20 degrees aoa up to about 30 degrees and at this aoa attain maybe 1.7 to 1.8. So, if you add a slat that spans the whole wing you can theoretically reach 1.7-1.8 at 30 degrees aoa. However, a part span slat will only help the part of the wing behind the slat. The rest will stall and loose lift at around 20 degrees aoa meaning the max Cl you can take out of the wing occurs at about 20 degrees, i.e around 1.2-1.5. So apparently it bears repeating: the slats on the Me109 are there to retain roll control at stall, not to improve Clmax. In fact, the Me109 wing does not incorporate a twist (wash out) like many other designs to retain roll control making the slats all the more important in this respect. I hope this makes it more clear how slats work and why they were included on the Me109.

 

So if you want to continue to push for a clean wing Clmax of 1.58 for the Fw190 and 1.7 for the Me109 in DCS it looks like you still have some ways to go.:smilewink:

 

Again not quite, as I already explained the inboard part of the wing stall at a higher AoA and achieves a higher CL max than the outboard part, thus adding slats to the outboard part WILL increase the overall CL max of the wing. And that was THE reason that Messerschmitt chose to use slats rather than washout.

 

A regular wing section will stall at anything from 15 to 24 deg, you cannot generalize and limit all wings to 20 degrees.

 

The CL max achieved using a similar wing planform as the 109 with the use of outboard slats again prove a figure of 1.7 - 1.8 as plausible.

 

That having been said you seem to have already made up your mind on this matter, irrespective of what'ever evidence is provided, so I don't expect you to change your opinion.

Edited August 31, 2014 by Hummingbird

[Hummingbird]

THe curves you posted have CL at AoA 0 about 0.6. It's absolutely impossible for the clean wings because, for example, that the plane must fly nose down attitude even at climb speed.

 

Typical CL_0 for the fighter plane is about 0.1 in clean configuration.

 

No, there is no 0 deg AoA CL figure, the line stops before 0 at around 2 deg AoA.