FDM: use reallife measured coefficients

[HHS81]

use real life measured coefficients from:
Brockhaus, R. (2013): Flugregelung. Physikalische Grundlagen. Mathematisches Flugzeugmodell. Auslegungskriterien-Regelungsstrukturen. Entwurf von Flugregelungssystemen. Entwicklungslinien. Springer Verlag Berlin Heidelberg GmbH

[xcvb85]

Good find 👍
Ich will check if I can get this book.

[HHS81]

Got the Ebook, unfortunately only readable via the browser. (but much, much cheaper - it is a 800page book...)

[HHS81]

With the input of the datas from the book, the aircraft behaves quite well, and feels really good.
Some datas were missing and had to be guessed.
One problem is that the Dornier 328 uses elevator trim tabs. Unfortunately the Dornier 328 needs 50% NU trim at takeoff. In reality only the yoke force changes, nothing else. In FlightGear the elevator output changes. Only ideas how to solve this with the new datas, but nothing to show yet.

[xcvb85]

Great 👍

[HHS81]

Ouch ...doesn't like stormy monday yet. 🌬

[xcvb85]

And doesn't like the new autopilot settings either. Let me know if I should adjust the autopilot.

[HHS81]

Rest now calm in windy, turbulent weather. Needs a lot of rudder in crosswind, and not much aileron. But 50ktn is too much - the real one has been tested up to 21ktn.

[xcvb85]

Please have a look at this:
https://books.google.de/books?id=tA7vCAAAQBAJ&pg=PA20

Too bad that Dornier prevented a publication of the drag scale but maybe it can be guessed:
https://books.google.de/books?id=tA7vCAAAQBAJ&pg=PA24

[xcvb85]

This is the lift/alpha curve from the wind tunnel test:

		<function name="aero/force/Lift_alpha">
			<description>Lift due to alpha</description>
			<product>
				<property>aero/qbar-psf</property>
				<property>metrics/Sw-sqft</property>
				<table>
					<independentVar lookup="row">aero/alpha-rad</independentVar>
					<tableData>
						-0.077 -0.30
						-0.057 -0.20
						-0.038 -0.10
						-0.018  0.00
						 0.000  0.10
						 0.020  0.20
						 0.037  0.30
						 0.056  0.40
						 0.076  0.50
						 0.095  0.61
						 0.114  0.71
						 0.132  0.80
						 0.152  0.90
						 0.172  0.99
						 0.191  1.07
						 0.210  1.15
						 0.230  1.22
						 0.249  1.29
						 0.267  1.35
						 0.285  1.29
						 0.303  1.27
						 0.321  1.30
						 0.338  1.31
						 0.357  1.25
					</tableData>
				</table>
			</product>
		</function>

[xcvb85]

I wonder why all of your new table values are depending on mach speed. It does not look right to me that drag decreases with increasing speed. Are you sure that you mean:

        <function name="aero/coefficient/CDalpha"><!--values are guess-->
            <description>Drag_due_to_alpha</description>
            <product>
                <property>aero/qbar-psf</property>
                <property>metrics/Sw-sqft</property>
                <property>aero/function/kCDge</property>
                  <property>aero/alpha-rad</property>
	<table>
                    <independentVar>velocities/mach</independentVar>
                    <tableData>
                          0.156	0.478 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                          0.303	0.072 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                          0.428	-0.002<!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                    </tableData>
                </table>
            </product>
        </function>

and not:

        <function name="aero/coefficient/CDalpha"><!--values are guess-->
            <description>Drag_due_to_alpha</description>
            <product>
                <property>aero/qbar-psf</property>
                <property>metrics/Sw-sqft</property>
                <property>aero/function/kCDge</property>
	<table>
                    <independentVar>aero/alpha-rad</independentVar>
                    <tableData>
                          0.156	0.478 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                          0.303	0.072 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                          0.428	-0.002<!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                    </tableData>
                </table>
            </product>
        </function>

[HHS81]

The dataset shows coefficients for three situations D1, D2 and D3 for a given weight of 10500kg:
D1: approach at mach 0.156, D2: climb at mach 0.303 and D3: cruise at mach 0.428.

Approach is specified with a flightpath of -3°, extracted gear and full flaps.
Climb with flightpath at 3° and no gear, no flaps and
cruise with 0°, no gear and no flaps.

Gear and flaps at approach explains the higher drag. I haven't included the flightpath in the table yet, but began today to filter out lift of flaps from the climb and cruise coefficients on may local system. I compared this new file with the source files mentioned in the ReadMe, especially the flightreport ""A better Pace setter" from flightGlobal and it comes nearly exact at the mentioned values.

So the values I set in are true and valid. The way the data is structured in approach, climb and cruise is quite common. The datas of c172p and c182s based on uses a very similar dataset.

I will send you the important pages of the books this weekend, unfortunately it is a Ebook, which is in a cloud. So I'm not able to copy the whole book 😞

[xcvb85]

OK, but then the high drag for low speeds is mainly caused by the flaps, gear and alpha, which is not considered here. So we must rather adjust the corresponding curves instead (cd_alpha, cd_gear, cd_flaps) to fit the specifications from your book.

[HHS81]

OK, but then the high drag for low speeds is mainly caused by the flaps, gear and alpha, which is not considered here.

That's what I said, and there are no explicit datas for the flaps and gear. (but for cd_alpha)
Deutsch: Die Widerstandwerte beinhalten bereits die Klappen und das Fahrwerk. Die spezifischen Koeffizienten dieser werden aber nicht genannt.

So we must rather adjust the corresponding curves instead (cd_alpha, cd_gear, cd_flaps) to fit the specifications from your book.

Unfortunately that's not that easy.
The dataset gives us the coefficients. With this we can build up the forces. This forces are all summed up at the end. (CLift = Clzero + Clalpha + Clalphadot + Cl....)
(as written by Gijs here)
But unfortunately we don't have the coefficients for the flaps and gear, and with that we don't have their forces.

We do know that they are burried in Clalpha and Clzero/ Cdalpha and Cdzero: CLift = Clzero (Clzero + Clflapszero) + Clalpha (Clalpha + Clflapsalpha) + .... And we do know at which settings (climb/ approach/ cruise) had been investigated. (bold= unknown)

So actually the coefficients of Clzero, Clalpha, Cdzero, Cdalpha etc... must be changed, so that we can create explicit forces for gear and flaps. Yesterday I think I was able to extract the flap coefficients from Clzero. But that's maybe not right yet.

So there are known variabales and unknown variables....
Deutsch: Das die Kräfte summiert werden, aber einige der Kräfte eben unbekannt sind, wäre es falsch die Koeffizeinten für Fahrwerk und Klappen zu ändern. Diese müssten irgendwie herausgefunden werden und aus den gegebenen Koeffizienten herausgerechnet werden.

[HHS81]

https://sourceforge.net/p/jsbsim/mailman/message/35749344/

[xcvb85]

Do you already have a CDalpha curve or should I extract it from the wind tunnel data?
The good thing is that CLzero is already included in the CLalpha data (at 0 alpha we have 0.1 lift).
If we find a meaningful value for CDgear we can extract CDflaps from your data. Maybe planes with similar gears (e.g. ATR42/72 or BAe 146) can help if we don't find anything for DO328.

[xcvb85]

Here you can see the value of CDgear:
https://books.google.de/books?id=axPvAgAAQBAJ&pg=PA538

[HHS81]

Do you already have a CDalpha curve or should I extract it from the wind tunnel data?

I have aero/coefficient/CDalpha due to mach, which is more than good enough.

Here you can see the value of CDgear:

It is just for the fairing, not the extracted gear.

Btw.:I tested your CLalpha curve, and used the CLzero to calculate CLFlaps. But the CLAlpha gives only values at 70m/sec, and tested in FlightGear shows me that the curve brings me not the correct AOA values I would expect, so I won't use it. I use CLalpha due to mach from the book, which gives me indeed the expected AOA values.

[HHS81]

Need someone who explains me how to deal with such curves, because I wonder why I don't get the AOA values who can expect....

[xcvb85]

I tested your CLalpha curve, and used the CLzero to calculate CLFlaps. But the CLAlpha gives only values at 70m/sec, and tested in FlightGear shows me that the curve brings me not the correct AOA values I would expect, so I won't use it. I use CLalpha due to mach from the book, which gives me indeed the expected AOA values.

??? AOA and alpha is the same and this is the input value. The lift force is the output value. Flaps shift this curve upwards (addition) and I guess the ground effect rotates this curve (multiplication).
I do not really understand why your values are depending on mach speed because the speed is already part of the pressure. The formula is: drag force = drag coefficient (cd in english or cw in german - also used for cars) * pressure * wing area

[HHS81]

??? AOA and alpha is the same and this is the input value. The lift force is the output value. Flaps shift this curve upwards (addition) and I guess the ground effect rotates this curve (multiplication).

I tested your curve instead of the coefficients from the book, and jsbsim/ FGFS gave me much higher AOA at certain speeds then it should. I'm sure the curve is correct, but it seems something I miss when it doesn't provide enough lift.

I do not really understand why your values are depending on mach speed because the speed is already part of the pressure. The formula is: drag force = drag coefficient (cd in english or cw in german - also used for cars) * pressure * wing area

First) MACH depends on temperature.
Second) the values from the book are real life measurements and not windtunnel. So the they don't have a calibrated weather. With using MACH they can be sure that the outputs will are always correct at any temperature.
Again: the dataset gives CL**/ CD** due to mach and not due to alpha, and that's how it works. Just look at the 747-200 which uses real measured datas....

[xcvb85]

What should the AOA values be? I get 2.7° at 3000ft altitude hold.

[HHS81]

What should the AOA values be? I get 2.7° at 3000ft altitude hold.

which indicated airspeed? which weight?

[HHS81]

Example: At 29000ft at 190ktn IAS the aircraft should have an Body Pitch Angle of 2.5° and with that an wing-aoa of 4° with an weight of about 27000lbs

[xcvb85]

I guess the wrong aoa has something to do with the pitch moments and not with the lift. Nevertheless thank you for the documents! Some of the values are a little strange in my opinion (e.g. cAalpha of 6.76 which is very high). I'm not sure how to proceed.

[HHS81]

Sorry for my late answer.
Distracted by real life and my buggy notebook.

Calpha 6.76 seems correct to me: The Dornier 328 uses the TNT wing of the Dornier 228, which had been designed to create much lift. Then the fueselage produces some extra lift (I read something in a old FlugRevue of my brother of 5% and more), and the Calpha of 6.76 is at approach with flaps full extracted.

Alan Teeder answered, see: (https://sourceforge.net/p/jsbsim/mailman/message/35749344/), but I'm not yet sure if I understood him correct. I think I have to subtract longitudinal (pitch drag lift) effect of flap and undercarriage from one to case to other. I have to ask him if I'm right how I have done it here on my local system, by changing those values.

[xcvb85]

From my understanding we should do the following for a nice FDM:

• Use the wind tunnel data as foundation (cl_alpha & cd_alpha)
• Adjust cl_flaps, cd_flaps, cd_gear to match the data from your book
• Adjust pitch_alpha to match correct AoA at different altitudes (where did you find this information?)

Do you agree or do you have a better idea? I know the problem is to distinguish between the different effects and I'm not sure if this will work.

[HHS81]

After digging a bit into aeroydynamics the last weeks I have to disagree 😞

Looking at the windtunnel data you found, I can see that study is only done to test those tripping devices and riblet films. The model used has no emepennage and no vertical and horizontal tail (p. 13), so the values are not realistic. The lift curve is only available for a speed of 70m/s = 136kts; and with tripping device. Drag polars are only available in a speed range of 50m/s to 90m/s = 97kts to 175kts.

The data from Brockhaus has a larger speed range from approach at 53m/s =102kts to climb at 103m/s= 198kts to cruise at 144m/s = 279kts.

I can also now answer why a lot of datas depends on mach speed. According to (NASA) above 250mph = 217kts mach compressibility comes into account:

("As an aircraft moves through the air, the air molecules near the aircraft are disturbed and move around the aircraft. If the aircraft passes at a low speed, typically less than 250 mph, the density of the air remains constant. But for higher speeds, some of the energy of the aircraft goes into compressing the air and locally changing the density of the air. This compressibility effect alters the amount of resulting force on the aircraft.")

The Dornier 328 has a much higher cruise speed than 250 mph, and the dataset takes mach compressibility into account.

The only way I can use the lift curve is to determine the stall angle together with the known stall speeds better.

[xcvb85]

The stuff they mounted on the wings had some effect on the drag but the influence on the lift was slim to none. The cl_alpha curve does more or less represent the quality of the airfoil and it is dimensionless. To get the lift force from the cl_alpha curve JSBsim multiplies it with the wing area and the pressure. The pressure is also calculated by JSBsim depending on the flight level and the mach speed so you don't need to care about this. Maybe the elevator has a slightly different cl_alpha curve but since the main lift is generated by the main airfoil this is not important. In my opinion the more input data we put into the FDM the better the result will be.

[HHS81]

When you disagree anyway, why do you ask me for my agreement?

Below Mach 0.3 cl_alpha is dimensionsless. Above it will be affected by mach compressibility.
From NASA:

To correctly use the lift coefficient, we must be sure that the viscosity and compressibility effects are the same between our measured case and the predicted case. Otherwise, the prediction will be inaccurate.

For very low speeds (< 200 mph) the compressibility effects are negligible. At higher speeds, it becomes important to match Mach numbers between the two cases. Mach number is the ratio of the velocity to the speed of sound. So it is completely incorrect to measure a lift coefficient at some low speed (say 200 mph) and apply that lift coefficient at twice the speed of sound (approximately 1,400 mph, Mach = 2.0). The compressibility of the air will alter the important physics between these two cases.

If that wouldn't be the case aircraft developing in real life would be very easy.....
No expensive flight tests, no complicated simulations, just simple school math.

But no, jsbsim does not compute mach compressibilty itself - it needs a table for, since every aircraft differs.

To quote Alan Teeter:

JSBSim does not directly apply changes to derivatives/coefficients due to
airspeed, Mach no, altitude or Reynolds number[sic!]. In other words all of these
must be done by the user - usually by interpolating between sets of data
tables. [...] In other words JSBSim is just a calculator, it is not a aerodynamics data
generator such as Datcom.[...]

I have just looked at the code and JSBSim does calculate the Reynolds
number, and ties the property. That saves one step.

He also wrote in the wiki:

You can write CLwing = CLo + Alpha* dCL/Alpha, or alternatively use a table to generate CL as a function of alpha, Mach no etc. It is not necessary to slavishly follow the examples in the JSBSim reference manual, or copy other Flightgear examples. It is best to follow the conventions that apply to the data you have available.

Believe it or not, but datasets from NASA for other aircraft gives all their values as due to mach. For some good reasons.....

I don't see why using this simply cl_alpha is better than cl_alpha due to mach

In my opinion the more input data we put into the FDM the better the result will be.

More garbage in, more garbage out.

[xcvb85]

ok

[xcvb85]

I made some tests with xflr5. Here are my input files:

a5.dat

DORNIER A-5 AIRFOIL
1 0 
0.98004 0.00828 
0.96008 0.01477 
0.94012 0.02156 
0.9001999 0.03313 
0.86028 0.04422 
0.82036 0.05309 
0.78044 0.06168 
0.74052 0.07006 
0.7006 0.07705 
0.68064 0.08024 
0.6646701 0.08243 
0.64072 0.08543 
0.62076 0.08782 
0.6008 0.09002 
0.58084 0.09182 
0.56088 0.09341 
0.54092 0.09481 
0.52096 0.09621 
0.501 0.09721 
0.48104 0.098 
0.46108 0.0986 
0.44112 0.0988 
0.42116 0.099 
0.4012 0.0986 
0.38124 0.0982 
0.36128 0.0978 
0.34132 0.09721 
0.32136 0.09641 
0.3014 0.09561 
0.28144 0.09461 
0.26148 0.09341 
0.22156 0.08982 
0.2016 0.08703 
0.18164 0.08403 
0.16168 0.08064 
0.14172 0.07685 
0.1018 0.06844 
0.09182 0.06547 
0.07186 0.05948 
0.0519 0.0517 
0.04192 0.04711 
0.03194 0.04152 
0.02196 0.03473 
0.01198 0.0259 
0 0 
0 0 
0.01198 -0.01737 
0.02196 -0.02256 
0.03194 -0.02615 
0.04192 -0.02902 
0.0519 -0.03154 
0.07186 -0.03525 
0.09182 -0.0388 
0.1018 -0.04032 
0.12176 -0.04311 
0.14172 -0.04571 
0.16168 -0.04771 
0.18164 -0.0499 
0.2016 -0.05209 
0.22156 -0.05409 
0.24152 -0.05589 
0.26148 -0.05768 
0.3014 -0.06048 
0.32136 -0.06168 
0.34132 -0.06247 
0.36128 -0.06327 
0.38124 -0.06347 
0.4012 -0.06347 
0.44112 -0.06227 
0.46108 -0.06108 
0.48104 -0.05988 
0.501 -0.05828 
0.52096 -0.05629 
0.54092 -0.05429 
0.56088 -0.0519 
0.58084 -0.0493 
0.6008 -0.04651 
0.62076 -0.04331 
0.64072 -0.04012 
0.6646701 -0.03593 
0.68064 -0.03273 
0.7006 -0.02874 
0.7205999 -0.02435 
0.74052 -0.01976 
0.76048 -0.01497 
0.78044 -0.01078 
0.8004 -0.00738 
0.82036 -0.00439 
0.84032 -0.0022 
0.86028 0 
0.88024 0.002 
0.9001999 0.003 
0.92015 0.0029 
0.94012 0.0026 
0.9606 0.0023 
0.98004 0.001 
1 0 

I got it from here: http://airfoiltools.com/airfoil/details?airfoil=doa5-il but the data was in a wrong order.

wing.xwimp

do320
0 2.56 0 1 0 13 5 1 -2 DORNIER/_/A-5/_/AIRFOIL DORNIER/_/A-5/_/AIRFOIL
4.07 2.56 0 1 0 13 5 1 -2 DORNIER/_/A-5/_/AIRFOIL DORNIER/_/A-5/_/AIRFOIL
9.4 1.38 0.6 1 0 13 9 1 -2 DORNIER/_/A-5/_/AIRFOIL DORNIER/_/A-5/_/AIRFOIL
10.48 0.29 1.6 1 0 13 5 1 0 DORNIER/_/A-5/_/AIRFOIL DORNIER/_/A-5/_/AIRFOIL

Maybe this is helpful.

[HHS81]

Sorry, no.

I had already tests done with JavaFoil, xflr5, Datcom+, the predicted curve by airfoiltoools and other software in the past years. Also has Michael Selig the coordinates of the airfoil as well.
None of the this brought me near to the outputs described in several reports.
There was another team working for a dornier 328jet for FlightGear, they used datcom as well, but the aircraft behaved also terrible wrong.

Again:

1. the wing alone doesn't help us, as it only provideds a fraction of the lift and drag. The fuselage also provides a whole lot due its unique shape.

2. according the FM the wing is blended of two different airfoil. The inner section uses the DO A-5M40 with 17% hickness, the outer section DO A-6M3 with 13% thickness. Only the coordinates of DO A-5 are known.

3. the reason why those programs Javafoil, xflr5 etc, above don't work for us as some engineers told me, is, that they regard it only as 2d- the wing itself is 3d. Which means it doesn't account the taper, twists, shape, blending of different airfoils of the wing - all things which also influences all those coefficients.

At least, to quote Alan Teeder, from the J3Cub-thread (https://forum.flightgear.org/viewtopic.php?f=4&t=28303&start=540#p315727), as they face a similar problem like we had:

Perhaps there are wind tunnel or flight test data publically available. These should give better results.
Anyone with a private pilots license and undergraduate aerodynamics training should be able to instrument an aircraft and conduct a simple flight test to get this parameter. Doing this was part of my degree course in the 1960´s.

And we do have have real data as provided by me above. And they do work and gives me finally the output I expected and are in line with those from the flight test reports. No need to struggle with predicted airfoils etc....

Real-life flight data > Wind Tunnel Tests >Datcom/ AVL > Aeromatic ! 😄

[HHS81]

JSBsim doesn't work good by only seeing the wing. JSBSim needs to see the whole aircraft to work correct.
YASim would work with the airfoil above, and it does, but has major flaws and missing features. YASim is beside Helicopters a No-Go for me.

[xcvb85]

In the end JDBSim just uses standard mechanical equations to put everything together and the question is what we should use for main lift and drag force (let's focus on this first because this is most important). Both forces are mainly depending on the AoA and air pressure (and wing area but this is constant) and currently I have no clue how you want to detemine these formulas.

[HHS81]

currently I have no clue how you want to detemine these formulas.

Das ist doch alles in den Coeffizienten von Brockhaus enthalten. Und die Tatsasche, dass das Flugzeug mit diesen Coeffizienten das macht, was es soll, zeigt mir, das diese viabel sind!

Die 747-200 hat auch keine klassische Liftcurve, wie Du es Dir vorstellst - und fliegt sich trotzdem nach Handbuch!

Das, was Du gerne hättest, zeigt nur die Extra 500.
Die Autoren haben aber Unterstützung von Extra Aircraft bekommen! Ich bin schon froh, das wir von den Ingenieuren der Dornier 328 Auszüge aus dem Manual bekommen haben.

[HHS81]

Vorschlag:
wir machen zwei fdms - eins auf Basis Deiner Methode, eins auf Basis meiner Methode.
Am Ende können wir es jeweils mit den Ergebnissen der Pilot/ Flight Test Reports etc. vergleichen.

[xcvb85]

Ich wollte nicht sagen, dass ich alles besser weiß, sondern nur verstehen, was Du konkret vorhast.
Meiner Meinung steckt im FDM der 747-200 ja doch eine klassische Lift-Kurve dahinter, da die Werte nochmal mit "aero/alpha-rad" multipliziert werden. Die Lift-Kurve ist in dem Fall eine Gerade, die durch den Ursprung geht. In Wirklichkeit hat die Kurve jedoch einen Offset (bei alpha=0 ist Auftrieb vorhanden) und das wird anscheinend über irgend einen anderen Effekt (Bodeneffekt?) zurechtgebogen. Komplettes Überziehen wird so nicht genau abgebildet (alpha > 10°), aber das finde ich auch nicht so wichtig.
Meiner Meinung nach ist der sinnvollste Weg, alles was man hat zu sammeln und dann zu schauen, wie man das zusammen bringen kann.

[xcvb85]

Ich war ein bisschen blind. Der Offset steckt natürlich in "Lift_at_zero_alpha". Das kann man natürlich so machen, wobei ich mir nicht ganz sicher bin, woher all diese Werte kommen.

[HHS81]

Meiner Meinung steckt im FDM der 747-200 ja doch eine klassische Lift-Kurve dahinter, da die Werte nochmal mit "aero/alpha-rad" multipliziert werden. Die Lift-Kurve ist in dem Fall eine Gerade, die durch den Ursprung geht. In Wirklichkeit hat die Kurve jedoch einen Offset (bei alpha=0 ist Auftrieb vorhanden) und das wird anscheinend über irgend einen anderen Effekt (Bodeneffekt?) zurechtgebogen.

Der "Offset" ist in der Tat Lift_at_zero_alpha. Das ist in der Aeroydynamic eine gängige Ausdrucksweise um eine Kurve zu beschreiben, die eben Nicht durch x= 0 geht. Ich kenne das auch von den Airfoils, die ich bei den Helis verwendet habe.

Das Überziehen wird bei der 747-200 nicht genau abgebildet, da dafür offenbar die Werte/ Daten fehlen. Die Daten wurden ja in Zeiten kreiert, wo solche Extremfälle rechnerisch nur schwer zu erfassen waren.

[HHS81]

Aber auch in dem fdm, dass ich mit den Bockhaus-Coeffizienten erstellt habe, steckt eine klassische Lift-Curve dahinter: https://github.com/HHS81/do328/blob/Issue69/do328-300.xml

<function name="aero/coefficient/CLalpha"> <!--                    <property>aero/alpha-rad</property>-->
                <description>Lift_due_to_alpha</description>
                <product>
                    <property>aero/qbar-psf</property>
                    <property>metrics/Sw-sqft</property>
		<property>aero/alpha-rad</property>
		<table>
                        <independentVar>velocities/mach</independentVar>
                        <tableData>
			0.0		0.0
                              0.156	6.76 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                              0.303	5.98 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                              0.428	6.00<!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                        </tableData>
                    </table>
                </product>
</function>

Auch hier wird mit alpha-rad multipliziert. Der Unterschied zu 747-200 ist, dass ich den Bodeneffekt nicht eingefügt habe. Der ist bei allen JSBSim-aircrafts ohnehin nur generic und damit eigentlich falsch 😆 .

[HHS81]

Man kann aber aus dem CLift-Coeffizienten, CLzero-Coeffizienten und alpha-rad auch eine Tabelle a la c172p erstellen.
Ungefähr so:

            <function name="aero/coefficient/CLalpha">
                <description>Lift_due_to_alpha, dependant on Mach</description>
                <product>
                    <property>aero/qbar-psf</property>
                    <property>metrics/Sw-sqft</property>
                    <property>aero/function/kCLge</property>
                    <table>
                        <independentVar lookup="row">aero/alpha-rad</independentVar>
                        <independentVar lookup="column">velocities/mach</independentVar>
                        <tableData>
                                      0.00    0.3
                             -0.09   -0.6084   -0.5382
                             ....
                              
                          </tableData>
                    </table>
		</product>
            </function>

Indem man für sich alpha-rad mit Brockhaus Cl-Coeffizienten multipliziert, Clzero dann dazuaddiert und das ganze in wie oben als Tabellenform darstellt.

Es bleibt die Herausforderung, Clflaps etc. vorher rauszurechnen. Wie, hat ja Alan Teeder ja auf der JSBSim-Mailinglist ja beschrieben und wird meine Aufgabe, sobald ich das Cockpit fertig texturiert habe.

[xcvb85]

Aber auch in dem fdm, dass ich mit den Bockhaus-Coeffizienten erstellt habe, steckt eine klassische Lift-Curve dahinter: https://github.com/HHS81/do328/blob/Issue69/do328-300.xml

Ist das der Auftrieb inklusive Landeklappen? Wenn ja, dann muss dieser Effekt da noch rausgerechnet werden, oder? Oder stecken die Landeklappen alleine in CA0? Ich bin mir auch nicht ganz sicher, was "Steigflug" bedeutet. Sind da noch Landeklappen 12° gesetzt? Und bedeutet CAalpha wirklich, dass diese Werte mit Alpha-Rad multipliziert den richtigen Auftrieb geben? Dann sind die 6.76 ja doch nicht so viel und entspricht in etwa dem, was ich für realistisch halte.
Beim FDM in Master bin der Meinung, dass der Auftrieb der Landeklappen irgendwie zu niedrig und der Auftrieb der Tragflächen zu groß ist. So kann man auch komplett ohne Landeklappen mit Vref landen (wenn auch mit sehr wenig Spielraum, z.B. Vref 110kt, Strömungsabriss bei 107kt). Hast du vielleicht ein Dokument, in dem die Stall-Geschwindigkeiten drinstehen?

[HHS81]

Ist das der Auftrieb inklusive Landeklappen? Wenn ja, dann muss dieser Effekt da noch rausgerechnet werden, oder?

Ja, inklusive Landerklappen, aber nur beim Approach.
Ja, Effekt muss rausgerechnet werden- Alan Teeder hat mir beschrieben, wie wir das am besten machen.

Oder stecken die Landeklappen alleine in CA0?

Nein

Ich bin mir auch nicht ganz sicher, was "Steigflug" bedeutet. Sind da noch Landeklappen 12° gesetzt?

Nein.
Auch nicht im realen Flug - da werden Sie frühstens bei 400ft AGL, spätestens bei 1000ft eingefahren. Je nach Fluggesellschaft.

Und bedeutet CAalpha wirklich, dass diese Werte mit Alpha-Rad multipliziert den richtigen Auftrieb geben? Dann sind die 6.76 ja doch nicht so viel und entspricht in etwa dem, was ich für realistisch halte.

Beides Ja, wie ich ja sagte.

So kann man auch komplett ohne Landeklappen mit Vref landen (wenn auch mit sehr wenig Spielraum, z.B. Vref 110kt, Strömungsabriss bei 107kt). Hast du vielleicht ein Dokument, in dem die Stall-Geschwindigkeiten drinstehen?

Ich habe noch keine Stalls eingebaut, darum ist es so einfach. Aber deshalb bin ich ja so sehr hinter realen Daten hinterher.
Ich habe im fdm selber alle verfügbaren Daten und Zahlen genannt.

Stall speeds:
Stall speed @max weight clean: 113kt
" " " " flap 32 degrees: 91kt
Ansonsten: Vref= 1,3* Vs (Vs = Stall speed landing config) Dazu schicke ich Dir per Mail noch ein Dokument. Geht leider nur nicht-öffentlich....

So, und wenn man das alles berücksichtigt und gegen die Daten aus dem Artikel "A better Pace setter" fliegt (abgesehen vom Treibstoffverbrauch - muss das noch korrigieren), passt das alles erstaunlich gut!

[xcvb85]

So könnte man die Lift-Kurve aus dem Windtunnel in das FDM einbringen:

        <function name="aero/function/corrAlpha">
            <description>corrected_alpha</description>
            <table>
                <independentVar>aero/alpha-rad</independentVar>
                <tableData> <!--Source: Recent Developments in Turbulence Management, p. 20-->
                   -0.077   -0.080
                   -0.057   -0.060
                   -0.038   -0.040
                   -0.018   -0.020
                    0.000    0.000
                    0.020    0.020
                    0.037    0.040
                    0.056    0.060
                    0.076    0.080
                    0.095    0.102
                    0.114    0.122
                    0.132    0.140
                    0.152    0.160
                    0.172    0.178
                    0.191    0.194
                    0.210    0.210
                    0.230    0.224
                    0.249    0.238
                    0.267    0.250
                    0.285    0.238
                    0.303    0.234
                    0.321    0.240
                    0.338    0.242
                    0.357    0.230
                </tableData>
            </table>
        </function>
	    
        <function name="aero/coefficient/CLalpha">
            <description>Lift_due_to_alpha</description>
            <product>
                <property>aero/qbar-psf</property>
                <property>metrics/Sw-sqft</property>
		<property>aero/function/corrAlpha</property>
		<table>
                    <independentVar>velocities/mach</independentVar>
                    <tableData>
                        0.0     0.0
                        0.156   6.76 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                        0.303   5.98 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                        0.428   6.00 <!--Source: Brockhaus.R (2013): Flugregelung, p. 745-->
                        </tableData>
                </table>
            </product>
        </function>

Auf diese Weise hätte man noch das Überziehen drin. Ansonsten habe noch das gefunden:

        <function name="aero/coefficient/CDi">
            <description>Induced_drag</description>
            <product>
                <property>aero/qbar-psf</property>
                <property>metrics/Sw-sqft</property>
                <property>aero/cl-squared</property>
                <value>0.0308</value> <!--Source: Aerodynamic Design of Transport Aircraft, p. 542-->
            </product>
        </function>

[xcvb85]

OK, die Lift-Kurve aus dem Windtunnel macht keinen Sinn. Mehr als 10° Alpha kann man eh nie erreichen.
Stichwort: Deep Stall

[HHS81]

Was bedeutet dieses corrected-alpha und warum ist dieses notwendig?

[xcvb85]

Meine Idee war es, doch die Lift-Kurve aus dem Windkanal zu nehmen. Daher habe ich sie entsprechend skaliert, sodass man sie einfach durch alpha-rad ersetzten kann. Allerdings tritt ein überziehen erst bei Alpha >15° auf, was man nie erreichen kann. Daher macht das keinen Sinn.
Der Teil mit Induced_drag macht aber Sinn (Oswaldfaktor ca. 0.8).

[xcvb85]

Wobei sich leider das Buch "Aerodynamic Design of Transport Aircraft" auch widerspricht. Auf Seite 538 ist von e=0.94 die Rede, was bei einer Flügelstreckung von 11 einem Induced_drag von 0.0308 entspricht. Auf Seite 542 ist der Induced_drag allerdings 0.0361. Ich gehe eher davon aus, dass die 0.0308 stimmen und man sich beim Eintragen in das Diagramm vertan hat, wobei ich mir nicht sicher bin.

[xcvb85]

Die 0.0308 werden richtig sein. Ich vermute, dass man sich beim Eintragen in das Diagramm in der Zeile vertan hatte, da man nicht davon ausging, dass der induzierte Luftwiderstand der DO-328 so viel geringer als der von der Saab 340 ist. Dabei hat die Saab keine Winglets und im Gegensatz zur Dornier auch keine abgeschrägten Flügelenden.
Die korrigierte Liftkurve macht doch Sinn. So kann man auch Deep Stall simulieren.