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Dora stall speed


Crumpp

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Several months ago there was a discussion on the CLmax of the FW-190.

 

http://forums.eagle.ru/showpost.php?p=2368020&postcount=83

 

http://forums.eagle.ru/showthread.php?t=136596&page=9

 

This not a "call to action" to have DCS make any changes that I am aware of to the Dora. It is simply to fulfill my promise of posting Focke Wulf GmbH determination of the CLmax of the aircraft.

 

First of all, let's talk about how engineers determine CL max of an airfoil. Simply put, the 2 dimensional cross section of the airfoil gives us our CL max. When we mathematically turn that 2D airfoil into a 3 dimensional wing, the CL max remains the same but the angle of attack data is shifted as a function of drag due to lift production.

 

In the FW-190 we find two different airfoils used and some aerodynamic twist. The airfoils selected are the NACA 5 digit series both of which were popular with designers of the day. In fact, Grumman used both of these airfoils or variations of in their fighters designs in World War II.

 

http://m-selig.ae.illinois.edu/ads/aircraft.html

 

These airfoils have two different purposes. The root airfoil, NACA 2015.3 determines the stall point or when the aircraft reaches Vs and the wing no longer produces enough lift required to keep the aircraft in normal flight.

 

The wingtip airfoil, NACA 23009 modifies the stall characteristics leaving the pilot with some lateral control and softening the stall characteristics ensuring the wingtips remain in flight so that the airplane remains controllable and does not posses a dangerous stall.

 

We will examine the NACA 2D data for both of the airfoils and gain some insight into what it is telling us and range of possible Clmax's each can produce.

 

First we need to understand the concept of Reynolds number (RN). Simply put, RN is a measurement of the "stickiness" of the air.

 

RN is used primarily as a scaling factor. It allows the engineer to make a tiny sized model and then predict how the full sized aircraft will act.

 

RN = (Velocity* Chord)/ kinematic viscosity of air

 

Rearranging the basic RN formula:

 

Velocity = (RN*kinematic viscosity of air)/ Chord

 

Now, we do not have all the information we need to make an exact velocity determination because we lack the data required to make a good wind tunnel correction for the conditions the data was determined. That is ok. We are not looking for specifics we just need an idea of what is plausible. We can tell enough to say the FW-190 cannot achieve this RN or this RN is probably well above stall speed of the aircraft to see the range of coefficients of lift the airfoil can produce. It gives a sense of the plausibility Focke Wulf's determination of CLmax.

 

In other-words, is the CLmax used by Focke Wulf reasonable and since they are the experts in their own design....correct. Of course it is but some readers will never be convinced of that fact.

 

Here is the root airfoil data. Granted, it is not the 15.3% chord of the FW-190 but it is close enough to gauge possibility. We need specifics, Focke Wulf GmbH had those specifics and already did that legwork for us.

 

Immediately, and engineer will notice the shape of the polar. The abrupt loss of lift coefficient at the stall tells us this airfoil has a sharp and violent stall.

 

312wtv8.jpg

 

If we run the math on the RN and ballpark the scale it to find the velocity the FW-190 would have to achieve to achieve the CLmax range of the 2D airfoil we find:

 

Velcoity = (8900000*.000156927)/2 = 698 fps or 475mph

 

At an RN of 8900000, our 24 inch airfoil has to be traveling at ~475mph to achieve a Clmax of 1.7.

 

The FW-190 with its wing chord of 5.95 ft would have to travel at 234 mph. That is well above stall speed. Windtunnel corrections change that speed specific speed.

 

Will call that our high speed realm.

 

The next RN is 2600000 and delivers a CLmax of 1.5.

 

Velcoity = (2600000*.000156927)/2 = 408 fps or 139 mph.

 

Our speeds work out to 139 mph for our 24 inch airfoil section and 46 mph for the FW-190 wing at an RN of 2600000.

 

Well, the FW-190 cannot fly at 46 mph so that is our low speed value.

 

So we can say with certainty the 2D airfoil could produce a CLmax ranging from 1.5 to 1.7 and the 1.5 lower end is not representative of the FW-190 wing's CL max. The CLmax in all probability lies somewhere in the middle!

 

Now that we have our plausible range for the FW-190's CLmax, lets look at the angle of attack the root airfoil stalls.

 

The NACA 2015 series stalls at about 18 degrees. Note that as it will be very important later.

 

Now lets look at the tip airfoil, NACA 20009.

 

2a0lqw2.jpg

 

Right away, we can see the polar shows us this airfoil also has the characteristic violent stall of the NACA 230XX 5 digit series.

 

First the speeds were are looking at on the polar.

 

At an RN of 826000 our 30 inch airfoil needs to travel at 353mph and our FW190 wing at 217mph.

 

It is our high speed realm.

 

At an RN of 3850000 our 30 inch airfoils needs to travel at 163mph and our FW-190 wing at 69mph.

 

Well again, the FW-190 cannot fly at 69mph.

 

We will call that our low speed realm.

 

Putting it together, the NACA 2009 delivers a CL max range of 1.45 to 1.57 or so.

 

Looking at the polar we see the stall angle of attack occurs between 18.5 and 21 degrees.

 

So, the FW-190 wing is comprised of two airfoils which deliver a CLmax range of 1.5 to 1.7 at the root and 1.47 to 1.57 at the tip.

 

The CLmax of the wing will be the result of calculus based on each of the wing section airfoils.

 

Fortunately we do not have to guess or do the math based on incomplete knowledge of the design. Focke Wulf did that math for us and is it listed on the cut sheet used by the firms engineers.

 

j9uglc.jpg

 

The CLmax of the FW-190 series wing is 1.58. In examining the 2D data, their determination is not only plausible, it is most likely correct.

 

Caution is advised for transferring aerodynamic coefficients from one system to another especially for drag. However, because wing CL max is simply taken from the 2D airfoil data and represents our 1G power off CLmax, it has a good chance of transferring easily form one system to the next. Aircraft performance calculations are predicated on this fact.

 

A CLmax of 1.58 gives a 4270Kg FW-190D9 a 1G stall speed of 109mph.

 

It also gives excellent agreement with flight testing results in both Allied and Axis test's.

 

Now, lets take a look at the airfoils stall angle of attack data so see how Focke Wulf and Grumman put together two airfoils with very harsh stall characteristics into a wing with a gentle 1G stall.

 

The NACA 2015 series stalls at about 18 degrees on a fairly consistent basis. Looking at the NACA 23009 airfoil the stall angle of attack occurs between 18.5 and 21 degrees.

 

Focke Wulf used a common practice to add two degrees of twist to the wing. In principle this means the root airfoil will always be at 2 degrees angle of attack higher than our wing tip airfoil. So when the root stalls at 18 degrees, the tip airfoil will only be at 16.5 degrees angle of attack. Since our root airfoil no longer allows the wing to produce enough lift to increase the angle of attack in 1G flight, our wing tip airfoil will never reach stall angle of attack.

 

Since only part of the wing is stalled, the entire wings stall characteristics are changed from the harsh stall of the 2D data into something exactly like this:

 

28uocwz.jpg

 

Here the entire aircraft was placed in a truss in a wind-tunnel in France at wind-speed of 20 mph and the polar constructed. It was not to measure CLmax but to gauge the stall characteristics of the aircraft.

 

Immediately, the engineers sees the gentle 1G stall characteristics of the wing.

 

It was done both with and without a propeller mounted as well as with power on.


Edited by Crumpp
Removed a simplification error
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There are two very strange points in your conclusions: the stall speed of 109 mph (174 kph) looks like a stall speed for a landing with FULL FLAPS.

20 mph = 36 kph = 10 m/s(!) - is not a typical speed in the wind tunnel. :) As far as I remember it was about 36 m/s that is very close to the Re of stall speed for the full-scale plane.

 

And 1.5 from the wind tunnel is very close to 1.58 in the German table.

 

The well know tendency for CL vs Re - is to grow with Re.


Edited by Yo-Yo

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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There are two very strange points in your conclusions: the stall speed of 109 mph (174 kph) looks like a stall speed for a landing with FULL FLAPS.

 

Unstick speed is 150-160kph IAS...

 

The wing has to be producing a CL of 2(+) with take off flaps to achieve that.

 

I will post some good data from the RAE, too.

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And 1.5 from the wind tunnel is very close to 1.58 in the German table.

 

9 PS is not the zero drag with propeller mounted at 180kph...it is more like the 9 m/s and 13 m/s I read in the report.

 

http://forums.eagle.ru/showpost.php?p=2323239&postcount=37

 

 

The well know tendency for CL vs Re - is to grow with Re.

 

 

Exactly, I just proved it to cut down on the "white noise" and to show readers the purpose of Re.


Edited by Crumpp
added the link to the polar with power on

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Good post Crumpp :)

 

Unfortunately I don't believe our ingame Fw190 features this CLmax figure. Either that or the P-51 features a windtunnel derived CLmax, i.e. assuming a completely smooth surface for laminar flow, something which was NOT possible in production aircraft, let alone in the field.

 

A factory fresh condition NACA 6 digit airfoil (i.e. std. roughness) produced a CLmax around 1.3.

 

By comparison the NACA 23xxx series didn't suffer any negative effects to its Cd or CLmax characteristics under operational surface conditions, as one can observe in NACA tests concerning this specific subject = surface roughness effects on various airfoils.


Edited by Hummingbird
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Btw, as an interesting side note one can also observe on the FW tables how the high aspect ratio of the Ta-152H's wing raises the CLmax from 1.58 to 1.7. Coupled with the large reduction in Cdi this results in a highly efficient wing in terms of lift to drag ratio, and thus it is little wonder why this aircraft was considered such a fantastic turning aircraft.

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Let's look at what our CLmax must be in a know configuration of 13 degrees take off flaps.

 

Unstick speed is listed in IAS as 150-160kph for the FW-190A5 and FW-190A6 handbuch.

 

2uzqn2h.jpg

 

Weight for the FW-190A5 because it is the lightest and will return the most conservative CL range:

 

256bnu1.jpg

 

Some simple math in the BGS system to see the range of possibility for the Coefficient of Lift required in the FW-190A5 to lift off with 10-13 degrees of flap:

 

Sea Level on a standard day.

 

IAS = EAS (we won't both with IAS to CAS because the PEC curve is diminishing CAS is ~6-9mph slower than IAS. It does not matter for this case because slower speeds = higher CL required.)

 

Weight = 4106Kg = 9052lbs

 

S = 197ft^2

 

q = V^2/295 (295 is a correction factor for using Knots in BGS...it works and gives good agreement and is easy to follow)

 

150kph = 80.9KEAS

 

q = 80.9KEAS^2/295 = 22.186p/ft^2

 

CL = Weight/(q*S) = 9052lbs/(22.186p/ft^2*197ft^2) = 2.071

 

160kph = 86.4KEAS

 

q = 86.4KEAS^2/295 = 25.3p/ft^2

 

CL = Weight/(q*S) = 9052lbs/(25.3p/ft^2*197ft^2) = 1.81

 

Unstick speed or Vmu is the minimum safe lift off speed. It is forward (faster) than stall speed in the take off configuration and represent the slowest safe speed the aircraft will begin to fly.

 

Adding landing flaps (40 degrees) MUST increase the CL of the wing. It is not going to get smaller with landing flaps.

 

The Coefficients are easily attainable and closer to what is expected with the split flap design.

 

It can be said with certainty that a CL of 1.58 cannot represent the CLmax of the wing with Landing Flaps deployed.

 

What is the stall speed then of a fully loaded FW-190D9 if our wings CLmax is 1.58 as Focke Wulf says in their cut sheet.

 

4270Kg = 9413lbs

 

q = 94.4KEAS^2/295 = 30.2p/ft^2

 

CL = Weight/(q*S) = 9413lbs/(30.2p/ft^2*197ft^2) = 1.58

 

94.4KEAS * 1.15 = 109mph EAS = 175Kph EAS

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A factory fresh condition NACA 6 digit airfoil (i.e. std. roughness) produced a CLmax around 1.3.

 

FWIW...

 

Standard roughness is misleading. To be at the NACA standard roughness, the aircraft would have to be finished in sandpaper.

 

If the airplane can fly with "standard roughness"....the company won't be getting sued no matter how much the finish deteriorates!

 

For aircraft performance and design, the airfoils are considered smooth.

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FWIW...

 

Standard roughness is misleading. To be at the NACA standard roughness, the aircraft would have to be finished in sandpaper.

 

If the airplane can fly with "standard roughness"....the company won't be getting sued no matter how much the finish deteriorates!

 

For aircraft performance and design, the airfoils are considered smooth.

 

Problem was that even small bulges would ruin the laminar flow airfoils' windtunnel characteristics, where'as this didn't affect the NACA 23xxx series.

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  • ED Team
Let's look at what our CLmax must be in a know configuration of 13 degrees take off flaps.

 

Unstick speed is listed in IAS as 150-160kph for the FW-190A5 and FW-190A6 handbuch.

 

2uzqn2h.jpg

 

Weight for the FW-190A5 because it is the lightest and will return the most conservative CL range:

 

256bnu1.jpg

 

Some simple math in the BGS system to see the range of possibility for the Coefficient of Lift required in the FW-190A5 to lift off with 10-13 degrees of flap:

 

Sea Level on a standard day.

 

IAS = EAS (we won't both with IAS to CAS because the PEC curve is diminishing CAS is ~6-9mph slower than IAS. It does not matter for this case because slower speeds = higher CL required.)

 

Weight = 4106Kg = 9052lbs

 

S = 197ft^2

 

q = V^2/295 (295 is a correction factor for using Knots in BGS...it works and gives good agreement and is easy to follow)

 

150kph = 80.9KEAS

 

q = 80.9KEAS^2/295 = 22.186p/ft^2

 

CL = Weight/(q*S) = 9052lbs/(22.186p/ft^2*197ft^2) = 2.071

 

160kph = 86.4KEAS

 

q = 86.4KEAS^2/295 = 25.3p/ft^2

 

CL = Weight/(q*S) = 9052lbs/(25.3p/ft^2*197ft^2) = 1.81

 

Unstick speed or Vmu is the minimum safe lift off speed. It is forward (faster) than stall speed in the take off configuration and represent the slowest safe speed the aircraft will begin to fly.

 

Adding landing flaps (40 degrees) MUST increase the CL of the wing. It is not going to get smaller with landing flaps.

 

The Coefficients are easily attainable and closer to what is expected with the split flap design.

 

It can be said with certainty that a CL of 1.58 cannot represent the CLmax of the wing with Landing Flaps deployed.

 

What is the stall speed then of a fully loaded FW-190D9 if our wings CLmax is 1.58 as Focke Wulf says in their cut sheet.

 

4270Kg = 9413lbs

 

q = 94.4KEAS^2/295 = 30.2p/ft^2

 

CL = Weight/(q*S) = 9413lbs/(30.2p/ft^2*197ft^2) = 1.58

 

94.4KEAS * 1.15 = 109mph EAS = 175Kph EAS

 

THis lift off speed does not seem very good to calculate CL because of two factors - ground effect and TO power. The more accurate result would give stall speed in clean configuration (power-off for sure) or touchdown speed (stilll with ground effect though).

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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  • ED Team

http://www.wwiiaircraftperformance.org/fw190/Fw_190_Eng-47-1658-D.pdf

 

118 mph is mentioned for the clean stall =190 kph. It gives CLmax =1.26 for 4100 kg. THat is very close (regarding Re changes) to the wind tunnel tests.

 

And 168 kph for landing configuration.


Edited by Yo-Yo

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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THis lift off speed does not seem very good to calculate CL because of two factors - ground effect and TO power. The more accurate result would give stall speed in clean configuration (power-off for sure) or touchdown speed (stilll with ground effect though).

 

I agree it is not good for specifically nailing down the CL max but that is not the intention.

 

It is good for a ballpark figure and the 1.58 is unlikely to be the landing flaps CLmax. The coefficient of lift returned by using Vmu certainly agrees well with a split flap equipped wing.

 

Unfortunately, Focke Wulf did not care to test the stall speed AFAIK.

 

Here is the landing reference chart used by FW-190A8 pilots. By convention, Vref approximates 1.3 the 1G stall speed for landing configuration.

 

15ekdvo.png

 

Using that convention, the CL max returned throughout the curve is ~2.67 with 40 degrees of flap.

 

That ballpark seems to give agreement with the findings of the NACA investigation into trailing edge high lift devices.

 

6ixhxx.png

Summary of section data on trailing edge high lift devices.pdf

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http://www.wwiiaircraftperformance.org/fw190/Fw_190_Eng-47-1658-D.pdf

 

118 mph is mentioned for the clean stall =190 kph. It gives CLmax =1.26 for 4100 kg. THat is very close (regarding Re changes) to the wind tunnel tests.

 

And 168 kph for landing configuration.

 

:thumbup:

 

It is a small world, LOL.

 

That is the report I was going to post to show the 1.58 as 1G stall speed.

 

Basically, the report is uncorrected and without weight and balance information.

 

He uses an approach speed of 130 IAS which is fast by Focke Wulf data but not out of the realm of possibility in practical piloting.

 

The IAS data for stall, clean and in landing configuration is without any altitude data. The 94.4KEAS is equivalent airspeed.

 

It is very doubtful that pilot tested the stall speed at an altitude below 10,000 nor is that data corrected to standard. Most likely, it is just we he read on the airspeed indicator in the winter of 1943. If he climbed to 20,000 feet, account for colder than standard day effects on IAS and you convert the IAS to EAS.....the 1.58 becomes very plausible for the 1G clean configuration CLmax.

 

Like I said, the most damning evidence is the combination of wing airfoils and split flaps would have to been extremely poorly designed to only achieve a 1.58 CLmax with 40 degrees of flap.


Edited by Crumpp

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Oh yeah, I used an FW-190A4 ladeplan minus the ammo and then determined if he climbed what fuel he would burn and if he had the time to do the test.

 

It was tight, but it worked out.

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BTW,

 

Yo-Yo this thread is not at all a call for you to change anything. Our methods might be different but the general result is the same and gives good agreement.

 

http://forums.eagle.ru/showpost.php?p=2369933&postcount=105

 

http://forums.eagle.ru/showpost.php?p=2368020&postcount=83

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Yo-Yo says:

118 mph is mentioned for the clean stall =190 kph.

 

94.4KEAS * 1.15 = 108.56 mph EAS

 

108.56 mph EAS - (+1) = 109.56mph CAS

 

109.56mph CAS + 7 mph PEC = 116.56mph IAS = 117 mph IAS

 

That is not even an 1% error.

 

 

Now if we look just above the CLmax on Weiderstandaten Von Flugzuegen we can see Focke Wulf lists the Coefficients of Lift for:

 

CaR = cruise flight

 

CaA = Approach

 

CaSt = Climb

 

CaA (Approach) Clmax exactly fits this curve of approach speeds.

 

An 8800 lbs Aircraft in Landing configuration at a Vref of 170kph needs:

 

170kph = 106mph IAS +3mph PEC = 109 mph CAS = 109mph EAS * 1.15 = 94.7KEAS

 

q = 94.7KEAS^2/295 = 30.4p/ft^2

 

CL = Weight/(q*S) = 9800lbs/(30.4p/ft^2*197ft^2) = 1.47

 

Well that agrees with Focke Wulf's approach Coefficient of lift.

 

If our CL max in landing condition is 1.58 then we only need to reduce velocity to 91.4KEAS.

 

That is 3.3 knots between approach speed and the stall. A 7 knot gust and you are stalling on approach.

 

That seems like some very tight landing margins and not very likely.

 

 

http://i61.tinypic.com/15ekdvo.png

 

Of course you know changing Vref is typical of aircraft that have a large weight variance. In the Airlines, we "get the numbers" every take off and landing.

 

I thought it was interesting that the FW-190 pilots also "got the numbers"!

EB-104correction.thumb.jpg.e2218a89dad8612ae9365ea78415d468.jpg

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Crumpp says:

94.4KEAS * 1.15 = 108.56 mph EAS

 

108.56 mph EAS - (+1) = 109.56mph CAS

 

109.56mph CAS + 7 mph PEC = 116.56mph IAS = 117 mph IAS

 

That is not even an 1% error.

 

If you work the Vs2 back to EAS from 105 mph IAS...

 

105mph IAS - 7 PEC = 98mph CAS - 1CeC = 97mph EAS = 84 KEAS

 

I chalked up his fast approach speed to performing an in-flight stall at altitude on a winter's day. He is doing the same thing I do when when I fly a new airplane. Take it up to up to altitude and stall it in landing configuration to get your approach speed.

 

105 IAS * 1.3 = 136 mph IAS which is why he used the 130mph IAS ballpark for Vref.

 

His Indicated Airspeed will read fast in the winter at altitude. It takes a 3% error to put us right back in the CL range of a split flap system.


Edited by Crumpp

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Good post Crumpp :)

 

Unfortunately I don't believe our ingame Fw190 features this CLmax figure. Either that or the P-51 features a windtunnel derived CLmax, i.e. assuming a completely smooth surface for laminar flow, something which was NOT possible in production aircraft, let alone in the field.

 

A factory fresh condition NACA 6 digit airfoil (i.e. std. roughness) produced a CLmax around 1.3.

 

By comparison the NACA 23xxx series didn't suffer any negative effects to its Cd or CLmax characteristics under operational surface conditions, as one can observe in NACA tests concerning this specific subject = surface roughness effects on various airfoils.

 

That is really Ok, Hummingbird. Just as there are different ways to skin a cat, there is different theory in aerodynamics. Mixing them can bring trouble. Even switching between the BGS and SI can deliver slightly different results.

 

What is important is not the specifics but the performance trend. There is margin of error in everything. The math shows us what is possible and simply a description of the physical world. What matters in the result in the physical world.

 

Outside of a small Rate of Turn difference, the analysis are almost the same.

 

You know, I have seen several folks post their math and theory pushing it off as some kind of holy grail of absolute truth.....that is not really true or correct. Yo-Yo models reflect reality "in the realm of significant digits" as one of my college professors was so found of saying. That is all you can ask of him.

 

As for the P-51...

 

I used the stall speed chart presented in the POH and interpolated for the max TO weight of 9611lbs for a stall speed of 102.8 mph IAS WITH wing racks.

 

I then corrected that IAS to CAS at sea level using a P-51 PEC curve.

 

102.8mph + 5.8mph = 108.6 mph CAS; CAS = EAS at sea level = 108.6 mph EAS*.869 = 94.3KEAS

 

An 9611lbs Aircraft traveling at 94.3KEAS needs:

 

q = 94.3KEAS^2/295 = 30.14p/ft^2

 

CL = Weight/(q*S) = 9611lbs/(30.14p/ft^2*235.75ft^2) = 1.35

P-51D_15342_Airpseed_Calibration.thumb.jpg.c5ac586988c1ac417d57f8455359bf7e.jpg

51ss.JPG.b2c6866e77d7e9032efbe295ed71ca4d.JPG

Limitations on P-51D and K.pdf


Edited by Crumpp
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  • ED Team

The next nail from my side :) :

 

You operate wind tunnel data. As well as German-French WT test. But the point is that the flight tests always shows lower CLmax because of trim losses - I mean stabiliser negative lift that is very sufficient at high trimmed CL's.

 

And the second point, I guess, you must take in account: the wing area taken for the plane generally includes the area of the fuselage that has very different lift/drag properties. The actual lift for the whole plane is produced by the wings (of less area that is specified for the plane) and the fuselage. So, the airfoil section numbers must be used very carefully.

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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So, the airfoil section numbers must be used very carefully.

 

In subsonic in-compressible flow theory as I was taught in college...the airfoil 2D data is used and converted to a 3D airfoil using the induced angle of attack.

 

I also use it because it was the predominate theory at the time in the 1940's with a few variations.

 

In the BGS system:

 

18.25 = correction factor for knots

 

Cl = coefficient of lift

 

AR = Aspect Ratio

 

Induced Angle of Attack = 18.25(Cl/AR)

 

3D wing angle of attack = 2D Angle of Attack + Induced Angle of Attack

 

So...in subsonic in-compressible flow theory you USE the 2D data. It is a key element to the determining the performance. For example drag computations, the reference area is designed to just use wing area...not wetted area. Use wetted area and it will deliver nonsense.

 

 

The next nail from my side :) :

 

You operate wind tunnel data. As well as German-French WT test. But the point is that the flight tests always shows lower CLmax because of trim losses - I mean stabiliser negative lift that is very sufficient at high trimmed CL's.

 

And the second point, I guess, you must take in account: the wing area taken for the plane generally includes the area of the fuselage that has very different lift/drag properties. The actual lift for the whole plane is produced by the wings (of less area that is specified for the plane) and the fuselage.

 

This is far down in the weeds.

 

If you do this with the system I am using...you will over think it and get the wrong performance. It is set up to use the 2D data and adjusted for conditions from that point.

 

See, that is why I was not screaming for you to change things.


Edited by Crumpp

Answers to most important questions ATC can ask that every pilot should memorize:

 

1. No, I do not have a pen. 2. Indicating 250

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That is really Ok, Hummingbird. Just as there are different ways to skin a cat, there is different theory in aerodynamics. Mixing them can bring trouble. Even switching between the BGS and SI can deliver slightly different results.

 

 

I just wish we would be using the actual real life figures, as that way we are sure that no bias is included.

 

The higher real life CLmax of the 190's wing is what allowed a much less powerful Fw190 Jabo to keep up with a light and souped up P-51B in a turning comparison during British tests despite a noticable disparity in wing loading and power.

 

All evidence available points towards a ~1.58 Clmax for the Fw190 and a ~1.35 Clmax for the P-51, that is when all factors are considerd. Thus I really did wish this was taken into account by Yo-Yo, esp. as I find it a lot better than relying on sketchy calculated data from wartime British tests which have since been known to be of a very dubious nature and wrong about a lot of things.


Edited by Hummingbird
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just wish we would be using the actual real life figures, as that way we are sure that no bias is included.

 

All the systems use real life figures and deliver performance that tells us what the airplane will do.

 

Yo-Yo is using a very detailed mathematical model. If it accounts for such things as fuselage lift then he is getting the same performance result without the higher CL. He is multiplying a smaller Coefficient of Lift by a larger reference area.

 

Think of it like this:

 

2*4 = 8

 

1*8= 8

 

Different numbers for the same answer.

 

 

That is why our performance agrees. The lines on a chart can be misleading too. It looks like a larger difference in Rate of Turn but it actually ~2 degrees a second. What both show is if either pilot tries to match the other...he will lose. The FW-190 cannot fly as slow as the P-51 and its ROT deteriorates if the pilot tries to match the P-51. Lag pursuit in the turn and keep your speed up. You will out-turn Mustangs in the Dora in DCS. If you look behind the P-51 and try to put your aircraft at that point and resist the urge to haul back on the sitck..it works as it should.

 

If you look AT the P51 and try to match its geometry and speed...you going to find your Focke Wulf full of .50 caliber holes.

 

The airplanes are equal dogfighters in both mathematical analysis and the game.

 

He can test it offline but if he is not using the same system or similar to that Focke Wulf used...the performance will not be correct.


Edited by Crumpp

Answers to most important questions ATC can ask that every pilot should memorize:

 

1. No, I do not have a pen. 2. Indicating 250

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  • ED Team

THis is Russian report on the 190.

 

Minimal speed (gauge reading) is 210 kph.

 

There is a landing speed observation WITH KNOWN MASS.

 

Летные испытания самолета FW-190 (ЭИ № 23 (219), декабрь 1943 г.)

 

Испытания проходил одномоторный трофейный истребитель FW-190A4 № 2310 с мотором BMW-801 выпуска 1942 г.

Сведения о самолете FW-190 приводились в ЭИ № 16 и 46, 1942 г. и в журнале «ТВФ» № 1, 1943 г. Здесь мы укажем лишь на особенности самолета и на отдельные агрегаты, хорошо показавшие себя во время летных испытаний.

По заключению НИИ ВВС КА, где самолет проходил испытания, детального изучения и внедрения на отечественных самолетах заслуживают:

1) компоновка винтомоторной группы в виде самостоятельного агрегата, включающего мотор, мотораму, маслобак, маслорадиатор, масляный фильтр и трубопроводы. Такое выделение ВМГ в отдельный агрегат очень выгодно в производственном и эксплуатационном отношениях;

2) принудительное охлаждение головок цилиндров и продувка маслорадиатора с помощью вентилятора, что обеспечивает нормальные температуры цилиндров и масла на всех режимах полета и работу мотора без применения юбки капота и заслонки маслорадиатора;

3) автоматизация управления винтомоторной группы, сосредоточенного в одном рычаге, что в значительной мере облегчает летчику ведение воздушного боя;

4) конструкция и система управления фонарем кабины;

5) электросистема убирания и выпуска шасси и посадочных щитков;

6) электрический синхронизатор пушки MG-151 и пулеметов;

7) патрон с электрозапалом для пушки MG-151, уменьшающий синхронное время;

8) управление подвижными щитками шасси, установленными на фюзеляже;

9) автоматический стопор костыля, стопорящий костыль при полном выбирании ручки управления на себя;

10) полная автоматизация и контроль работы стрелково-пушечного вооружения, что значительно облегчает работу летчика в бою.

Взлет

 

Положение щитков, град Полетный вес, кг Обороты мотора, об/мин Давление наддува, атм Длина разбега, м Скорость отрыва, км/ч Длина взлетной дистанции до набора Н = 25 м, м

0 3989 2450 1,35 520 165 1300

10 3989 2450 1,35 500 160 1250

 

Летная оценка самолета

 

 

Руление. Самолет легко управляется на рулении и устойчиво выдерживает прямолинейность ' направления. Наличие автоматического стопора хвостового колеса, связанного с ручкой управления (при ручке, взятой полностью на себя, хвостовое колесо стопорится), облегчает руление.

Ножные гидравлические тормоза эффективны.

Взлет. Особенность на взлете — большая величина пробега 520 м (без щитков), поэтому взлет, как правило, производится с выпущенными в стартовое (10°) положение щитками. Скорость отрыва при нормальном взлете равна 165 км/ч.

В начале разбега самолет имеет стремление к разворачиванию, что легко парируется педалями. Усилия на ручку для подъема хвоста значительны. Наилучшее положение стабилизатора для взлета +1,5 деления по указателю в кабине (что составляет 2").

Набор высоты. На наборе высоты после убирания шасси и щитков самолет тянет на нос.

Триммеры рулей высоты отсутствуют, и нагрузка на ручку снимается изменением угла установки стабилизатора; в связи с этим шасси следует убирать на скорости не менее 200 км/ч. Наивыгоднейшая скорость набора высоты до 5000 м составляет 275 км/ч и поддерживается до этой высоты постоянной. После 5000 м скорость набора следует уменьшать через каждые 1000 м на 5 км/ч. Вторая скорость нагнетателя включается автоматически при достижении высоты 2500 м; при этом наддув равен 1,35 атм.

Горизонтальный полет. Самолет устойчив во всем диапазоне скоростей. На максимальной скорости поведение самолета нормальное. Минимальная скорость 210 км/ч (по прибору).

 

Летные данные

 

 

Максимальная скорость на номинальной

мощности, n=2450 об/мин, p = 1,35 атм, км/ч:

у земли 510

на 1-й границе высотности H = 1800 м. 544

на 2-й границе высотности H = 6000 м. 610

Время подъема на 6000 м, мин 6,8

Практический потолок, м 10 500

Время набора практического потолка, мин. 32

Время выполнения виража, с.

на H = 1000 м 22

на H = 5000 м 30

Виражи левые посадочными щитками

Время выполнения боевого разворота при H = 1000 м, 21

Скорость ввода 500 км/ч,

скорость вывода 250 км/ч (по прибору)

Время переворота на H = 1000 м, с 15

Скорость ввода 270 км/ч, скорость вывода 450 км/ч

Время выполнения петли, с 21

Маневренность. На пилотаже самолет отличается большими переменными нагрузками на ручку управления рулем высоты.

Перекладывание из виража в вираж легкое. При выполнении виража со щитками, выпущенными в посадочное положение, самолет становится устойчивее и время виража уменьшается на 1–2 с. При выполнении вертикального маневра с набором высоты быстро теряется скорость.

Петля выполняется без потери высоты.

При выводе из фигур и из пикирования нагрузки на ручку чрезмерно большие. Для облегчения вывода необходимо пользоваться стабилизатором (ввиду отсутствия триммеров).

Пикирует самолет устойчиво, быстро набирая скорость.

Планирование и посадка. Самолет с убранными щитками устойчиво планирует на скорости 270 км/ч, а с выпущенными щитками — на скорости 240 км/ч (по прибору); эти скорости являются наивыгоднейшими для планирования на посадку.

Глиссада планирования крутая. При подходе к земле с полностью убранным газом на скорости 240 км/ч выдерживание над землей небольшое, самолет быстро теряет скорость и приземляется на три точки с почти подобранной ручкой. На пробеге и в момент приземления самолет устойчив. Длина пробега велика и равна 530 м.

Взлет и посадка производились на бетонированной дорожке. Посадка производилась с тормозами.

Полеты производились до полного выгорания горючего. Режимы максимальной дальности подобраны специальным полетом на определение расходов горючего по бензиномеру на различных скоростях полета.

 

Посадка

 

Положение щитков, град Полетный вес, кг Длина пробега,м Время пробега,с Посадочная скорость, км/ч Длина посадочной дистанции H = 25 м, м

60 3800 530 20,3 154 1120

Дальность и продолжительность полета

 

Высота полета, м Режим Скорость полета. км/ч Число оборотов мотора, об/мин Давление наддува, атм Дальность, км Продолжительность полета

5260 0,9 максимальной скорости 542 2100 1,1 552 1 ч 02 мин

1245 режим максимальной дальности 395 1700 0,97 983 2 ч 30 мин

Takeoff

 

Flaps, deg Mass, kg Engine rpm Manifold pressure, at TO dist, m TO speed, kph TO distance to obstacle Н = 25m, m

0 3989 2450 1,35 520 165 1300

10 3989 2450 1,35 500 160 1250

 

Landing

 

Flaps, deg Mass, kg Landing run, m Landing run time, s Landing speed, kph Landing distance from H = 25 m, m

60 3800 530 20,3 154 1120


Edited by Yo-Yo

Ніщо так сильно не ранить мозок, як уламки скла від розбитих рожевих окулярів

There is nothing so hurtful for the brain as splinters of broken rose-coloured spectacles.

Ничто так сильно не ранит мозг, как осколки стекла от разбитых розовых очков (С) Me

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