A discussion of how to correctly set up Corrected thrust table 1506

[jx_]

Go here for an explanation, http://fsdeveloper.com/forum/showpost.php?p=579052&postcount=24 then feel free to post questions in this thread.

UPDATE:

Here is a simpler explanation that may be easier to implement. http://fsdeveloper.com/forum/showpost.php?p=654582&postcount=5 It includes examples that weren't in the original thread.



Terms and formulas used in the discussion:

"Corrected" or "Normalized" - This means what one thing would be under different conditions. In aviation this is usually either converting from any flight condition to stationary, sea-level, in standard weather, or the opposite. In the explanation you will see corrected thrust and corrected n1 CN1. Corrected thrust is conversion from whatever condition you are experiencing --- to what thrust would be at the same throttle setting if we were stationary at sea level with standard weather. Corrected N1 is simply what n1 would be at sea level under standard conditions if we were producing the same thrust.

The formula to correct n1 is "Gauge n1 divided into the square root of Theta2."
The formula to correct thrust is "Current thrust in pounds divided into Delta2"

The formula to find what actual n1 would be from CN1 is "CN1 times square root of Theta2"
The formula to find what actual thrust would be from a static chart is "Thrust looked up on the chart times Delta2 for the conditions you are predicting."


Delta - Pressure ratio. As altitude in increased, air pressure decreases. Current pressure divided into standard pressure (29.92 or 1013mb) equals pressure ratio.

The formula for pressure ratio is "current local pressure / standard pressure" in same units. Here are the standard units:

2116.21662 pounds per square foot
14.6959 pounds per square inch
1013.25 millibars
29.92 inches of mercury


Theta - Temperature ratio. Same as pressure ratio, except current temperature divided into standard temperature (+15C or +59F converted to Kelvin or Rankine)

The formula for temperature ratio is

Temperature in Celsius + 273.15 / 288.15 or Temperature in Fahrenheit + 459.67 / 516.67




Delta2 & Theta2 - the 2 indicates total. This is sometimes referred to as Delta_t and Theta_t. The official designation is Total Pressure Ratio and Total Temperature Ratio. When something is total it is taking into account the effects of stagnation. Stagnation is when airflow is compressed as it comes into contact with an object moving at a different speed. As the surface of an airplane, bullet, car, etc push through the air, the air is slowed as pressure increases in front, then accelerated as pressure decreases around and behind (think of a boat making a pile of violent water in front and a vacuum that fills in behind as it speeds through the water.) In the high pressure area, the air becomes heated, changing its temperature. The change in temperature affects its density (pressure and temperature) characteristics which affect how the wings and engines convert this air into work. It is important because if the air entering the engine is hotter than the outside air temperature, the engine will perform exactly as if the airplane was stationary in hotter conditions.

The formula for Delta2 is (Delta * ( Mach * Mach * 0.2 + 1 )^3.5 power)
The formula for Theta2 is (Theta * ( Mach * Mach * 0.2 + 1 ))

 

[vololiberista]

That's very interesting jx. However, the 1506 table is not annotated. I have a top grey row and two white rows. In your explanation I cannot find what row/column serves what purpose. The figures you quote need to be shown diagramaticaly as well.
vololiberista

 

[jx_]

They are annotated in the information dialog (right click)

The Mach values are the far right grey, CN1 the top grey, and the white are the 1506 scalars of thrust.

 

[Heretic]

Hooray for necrophilia!

In both cases you MUST copy your static thrust to each mach line and convert to corrected RPM. 100%n1 at Mach 1.0 would be moved to 92.4% CN1. See note below.....

To set 1506 to net thrust is very simple:

Ram drag pounds / static thrust reference = scalar to subtract from MACH lines after you have corrected each mach line.

SO, 1000 Fr / 10000 SSL Fg = 0.10 Fr scalar

1506 scalar of 1.0 - 0.10 Fr scalar = 0.9 entered into 1506 at whatever CN1 vs Mach we were converting.

Can you provide an example table or a graph for this or something?
Is Fg gross thrust?
Also, how does table 1507 need to be set up for this case? Y=1 for every CN1 value given?


Compared to the other tables, this is bloody difficult. Despite having EngineSim, MATLAB and a pretty good set of internet based data on the JT8D(-17)*...



(*Compared to just about any other engine.)

 

[jx_]

Hello there,

If you will be using table 1507 (ram drag) as well as 1506 (gross thrust), then you only need one entry in table 1506. One entry for all mach with Sea Level static gross thrust. All altitude effects are taken care of internally and table 1507 will take care of ram drag at different mach.

But it is not necessary to use table 1507, as you can also use table 1506 as a corrected NET thrust table and leave 1507 all zeroes. This is in fact the easiest and quickest way to get accuracy if you have a corrected NET thrust table, which is much more common than a gross thrust table.

If you have corrected data on your JT8D, your work is easy.... zero out table 1507, then fill out table 1506 to match your data exactly. I have attached a real world plot for a JT9D that gives all the pertinent data you need to make an FS engine model that is accurate at any speed, and correctable to every altitude.


If you don't have such a plot then you need to find Ram Drag. There are a few ways to do this...


1) assuming intake losses are negligible: corrected mass flow (remember mass flow is RPM dependent and based on demand!!!) * intake area * True Airspeed in feet per second is equal to momentum drag. This method requires a mass airflow plot per CN1.

Fr = m.intake * Aintake * KTAS * 1.6878


2) Use sim or tool: This is the method I was trying to explain in the other post. After reviewing that post I believe I was unclear or confusing about some requirements and can explain it in much simpler terms...

- Set up your airfile with gross thrust in 1506 matching the real aircraft.
- zero all fields in table 1507
- grab cruise performance charts and load scenario
- after ensuring you have your drag correct, the FS gross thrust required to maintain the scenario speed is equal to net thrust
- correct the net thrust you found in the previous step
- correct the performance charted n1, then record cFN to that cN1 at that mach.

-OR-

- Set up your airfile with gross thrust in 1506 matching the real aircraft.
- Set up your airfile with ram drag in 1507 matching the real aircraft.
- grab cruise performance charts and load scenario
- correct the real world charted fuel flow, then divide into SSL TSFC
- correct the FS Net thrust
- after ensuring you have your total drag correct, the difference in thrust derived from actual corrected fuel flow and FS corrected net thrust is equal to Ram Drag for that sea level mach.
- correct the performance charted n1, then record cFN to that cN1 at that mach.

This is essentially; real world gross thrust - sim predicted net thrust = ram drag.

Always remember, ALL the numbers you see in table 1506 are sea level mach.

Both of these methods require accurate drag in the airfile.

The same steps can be done using engine sim or anything else to compare against the charts.

Do not take your FS corrected net thrust and apply it directly to static gross thrust. The goal is to find the difference the engine is experiencing at that mach. This is more accurate as it naturally includes all forms of ram drag.

cFF = Fuel Flow / ( Delta2 * Theta2^0.5 ) <-- 0.5 is square root, but this is not a constant. Some engines may vary from 0.4 to 0.6.

cN1 = N1 / Theta2^0.5 <-- again, the square root may vary for different engines.

TSFC = fuel flow / thrust reference


3) If you have a drag polar, or any other detailed drag plot, drag is simply thrust required. There is NO way to actually measure drag unless we put the full size airplane in a wind tunnel we can depressurize. So we find drag by using the net thrust plot. You can reverse this process. If you have a chart of CD (drag coefficients), you can match the CD to the snapshot to find the net thrust and plot it that way. You would solve for drag in pounds. At constant airspeed drag pounds is equal to net thrust.

Drag = CD * q * S,

q = 1481.4 * mach^2 * Delta

S = reference wing area, ft^2

Drag = Net Thrust


For a climb:

( (Thrust - Drag) / Weight ) * 101 * KTAS = climb rate in feet per minute, so...

( Feet per minute / ( KTAS * 101 ) ) * Weight = Excess Thrust

Excess Thrust + Drag = Net Thrust


And yes, Fg is the designator for Gross Thrust


I hope this is a simpler and understandable explanation. If not let me know and I will try again. The chart below is the real world equivalent of what the designers at Microsoft were intending to do when they made the airfile format the way it is. Why they didn't include TGT, EPR, mass flow, and N2 is beyond me. (well, table 1507 is a corrected mass flow table, but it is only used to find net thrust.)

Your goal here as an FD designer is to make a chart in your airfile similar to the one below. Those numbers you see below are not the actual numbers you will see at altitude, but corrected to sea level equivalent per CN1 at that mach. With this chart, we can uncorrect to any Delta2 and Theta2 combination. Those two variables are all you need!

It is an important concept to understand a JT9D will never see anything near mach 0.8 at sea level. However,, when the aircraft is aloft at mach 0.8, the engines are producing thrust equal to { sea level mach 0.8 * Total Pressure Ratio }


=============

According to the last line above, the JT9D engine will produce 14600 pounds of net thrust and burn 10200 pounds of fuel at 102%CN1, mach 0.8. These are corrected numbers that only apply to sea level (identical to what table 1506 uses). This is an impossible scenario and would destroy the engine at sea level, but at 25,000 feet:

Thrust = 14600 * Delta2 = 8273 lbs
Fuel Flow = 10200 * (Delta2 * Theta2^0.59) 5467 lbs
N1 = 102.4 * Theta2^0.5 = 98.96 %n1

FS will output these numbers the same way.

=============


And here is an example of application:

I've taken real world performance charts and corrected the n1 and fuel flow. Then found net thrust, corrected it and entered it above. Once you have enough plot points at different mach, you will be able to draw an arc that represents the entire Net thrust curve at each mach.

When you are done it may look like this: *this chart doesn't match the above table. I don't have any charts I can post publicly. But the concept is the identical.

 

[jx_]

Also, how does table 1507 need to be set up for this case? Y=1 for every CN1 value given?

I am going to assume you meant table 1505.

There is a post here for that as well, but the simple run down is

Y = CN2

X = Time

When looking at the table in aired, the steeper the red line, that faster the engine will reach the next CN2 (because the next X is closer.)


a flat line will result in an engine that will idle at that CN2 regardless of throttle position.

a vertical line will result in the engine instantaneously jumping from the lower CN2 to the higher CN2. It will always skip all CN2 in between because there is no X time between the two Y values.


FS complicates the above explanation because it has a bit of code to impersonate a fuel system, so for larger differences between commanded RPM vs current RPM, FS will shorten the time to make the engines respond quicker or vice versa for short distances. This appears to be some form of feedback equation.

jx

 

[Heretic]

Much more informative than the other post. Thanks!

I am going to assume you meant table 1505.

No, 1507. But you've covered it in the previous post.

 

[jx_]

Glad that helped

 

[Heretic]

I'm still confused.

Don't fear though, I already have a broad idea about this, but I need confirmation.

So...JT8D-17...

I've got my baseline N1 and thrust values for mach 0 and sea level (at ISA). Hence, they don't need to be corrected.
They were obtained from a table. A sparsely populated table, but a table nonetheless.

And this is the heart:
96.6% cN1 = 16000 lbs thrust

So this means, as long as static_thrust in the aircraft.cfg reads 16000, 1.0 for the thrust factor at cN1=96.6% for mach index 0 in table 1506.
Together with the other values from the table, I got a nice looking curve for mach index 0. So far, so easy, so good.

Now for the next mach index.

Since I'm at SL, so theta=1. I pick 0.45 as the next mach number.

Hence: theta2 = 1*(((0.45^2)*0.2)+1) = 1.0405

And thus: cN1 = N1/(theta2^0.5) = 94.70%

This means that 96.6% N1 at mach 0.45 (SL,ISA) corresponds to 94.7% N1 (SL,ISA) at mach 0, right?

So, applied to my fixed 96.6% cN1 base value, this means that at mach 0.45, I will get a thrust factor of slightly less than 1.0.
Lazy as I am, I would therefor just do something like:
(94.7/96.6)*[thrust factor at mach 0]=[thrust factor at mach 0.45]

Would this work?



And if I understood this correctly, in the big picture, MSFS reads the throttle lever state from the flight controller or autopilot, then runs the value through tables 1503 and 1504 (depending on mach and IAP) to obtain cN2, then applies 1502 to the result to obtain cN1 before pulling an appropriate "thrust factor" value from 1506 that is then multiplied with "static_thrust" from the aircraft.cfg to obtain the *uncorrected* thrust. Said uncorrected thrust is finally multiplied with "thrustspecificfuelconsumption" to obtain the current *uncorrected* fuel consumption and combined with some decent drag values, it makes your plane go as far as it's supposed to (which is my goal witzh this one).

 

[jx_]

And this is the heart:
96.6% cN1 = 16000 lbs thrust

So this means, as long as static_thrust in the aircraft.cfg reads 16000, 1.0 for the thrust factor at cN1=96.6% for mach index 0 in table 1506.

Yes, 1.0 = 16000 @ CN1 you entered 1.0.

Together with the other values from the table, I got a nice looking curve for mach index 0. So far, so easy, so good.

Now for the next mach index.

If you are using table 1507, you only need sea level gross thrust. If you are not using table 1507 and using table 1506 as a net thrust table, continue entering mach lines...

Since I'm at SL, so theta=1. I pick 0.45 as the next mach number.

Hence: theta2 = 1*(((0.45^2)*0.2)+1) = 1.0405

And thus: cN1 = N1/(theta2^0.5) = 94.70%

Correct.

This means that 96.6% N1 at mach 0.45 (SL,ISA) corresponds to 94.7% N1 (SL,ISA) at mach 0, right?

Yes. Now you are getting it.

(***Except for ram effect where net thrust increases with total pressure...FS handles that on it's own so we are not concerned with it here, but yes, the engine is indeed operating AS IF it was at the lower static N1.)

So, applied to my fixed 96.6% cN1 base value, this means that at mach 0.45, I will get a thrust factor of slightly less than 1.0.
Lazy as I am, I would therefor just do something like:
(94.7/96.6)*[thrust factor at mach 0]=[thrust factor at mach 0.45]

Would this work?

No. This would apply to the reverse. If you were trying to take actual thrust, and convert it back to plot corrected thrust, you would have to align that corrected thrust to corrected N1. When starting from corrected n1 and corrected thrust, you only need to keep the two values together.

This is the reason you don't need any mach lines if you are using table 1507. If you are making a net thrust table in 1506, the only difference between 96.6%cN1 at static and 96.6%cN1 at mach X will be a reduction for ram drag.

And if I understood this correctly, in the big picture, MSFS reads the throttle lever state from the flight controller or autopilot, then runs the value through tables 1503 and 1504 (depending on mach and IAP) to obtain cN2, then applies 1502 to the result to obtain cN1 before pulling an appropriate "thrust factor" value from 1506 that is then multiplied with "static_thrust" from the aircraft.cfg

EXACTLY correct up to there.

to obtain the *uncorrected* thrust.

Incorrect. The airfile tables are all corrected. This is why simply multiplying by the static_thrust line works, because we are still dimensionless. Up until this point above^^^ the sim thinks the engine is in ISA SL conditions.

Corrected means equal to ISA SSL
Uncorrected means at actual ambient.


I've edited your next statement in red...

Said corrected thrust is finally multiplied with "thrustspecificfuelconsumption" to obtain the current *corrected* fuel consumption

Then three things happen:

Thrust is uncorrected to ambient automatically by the sim: Corrected Thrust * Delta2 = Actual Thrust
Fan speed is uncorrected to ambient automatically by the sim: Corrected RPM * Theta2^0.5 = Actual N1
Fuel Flow is uncorrected to ambient automatically by the sim: Corrected Fuel Flow Weight * (Delta2 * Theta2^0.5)

You now have your primary engine outputs for any condition.

If they would have made an EGT chart it would have been just as simple, using the appropriate formula. Also, if using table 1507 for ram drag, this is how that works:

Mass Air Flow is uncorrected to ambient automatically by the sim: Corrected AirFlow Weight * (Delta2 / Theta2^0.5) = Actual Airflow

Actual Airflow * KTAS * 1.6878 = Total Ram Drag

and combined with some decent drag values, it makes your plane go as far as it's supposed to (which is my goal witzh this one).

yes. As far, with the correct amount of thrust at the correct fan RPM for the conditions. If the drag is exact, and the tables are exact, the sim will output EXACT!

(of course that's not really possible because real airplanes have defects and inconsistencies.)

Hope that helps. Sounds like you are on the right track.

jx

 

[Heretic]

No. This would apply to the reverse. If you were trying to take actual thrust, and convert it back to plot corrected thrust, you would have to align that corrected thrust to corrected N1. When starting from corrected n1 and corrected thrust, you only need to keep the two values together.

This is the reason you don't need any mach lines if you are using table 1507. If you are making a net thrust table in 1506, the only difference between 96.6%cN1 at static and 96.6%cN1 at mach X will be a reduction for ram drag.

So I take that my calculation should be the other way around?

(96.6/94.7)*[thrust factor at mach 0]=[thrust factor at mach 0.45]

This should be something like 1.02, so this would mean that I need 1.02 times more thrust at mach 0.45 to attain 96.6% N1 that I'd need at mach 0.

If this is still incorrect, what do I need to calculate (and in what way)?

(Note: I don't intend to use table 1507.)


Anyway, I really appreciate your input on this one. If I can pull this off with the 727, I might just write a short wiki-based "How To" for this case.




(P.S: If you want to bring your general knowledge about setting up 1502 to 1507 into the wiki...you can just start an article there without any kind of moderator approval. The only disadvantage is that there's no FDE section (yet).)

 

[jx_]

So I take that my calculation should be the other way around?

(96.6/94.7)*[thrust factor at mach 0]=[thrust factor at mach 0.45]

This should be something like 1.02, so this would mean that I need 1.02 times more thrust at mach 0.45 to attain 96.6% N1 that I'd need at mach 0.

If this is still incorrect, what do I need to calculate (and in what way)?

(Note: I don't intend to use table 1507.)


Anyway, I really appreciate your input on this one. If I can pull this off with the 727, I might just write a short wiki-based "How To" for this case.




(P.S: If you want to bring your general knowledge about setting up 1502 to 1507 into the wiki...you can just start an article there without any kind of moderator approval. The only disadvantage is that there's no FDE section (yet).)

No calculation to be done other than reducing gross thrust to net thrust.


There are two basics concepts to setting up thrust:

1) You have a thrust chart
2) You have aircraft data which needs to be converted to thrust required to find net thrust, then corrected thrust



In your case, I am under the impression you have a chart and have started with Gross Thrust. The only thing you need from here is to find net thrust at mach 0.x

Does your chart show net thrust at speed? You table 1506 should be gross thrust at mach 0.0, then net thrust at the other mach.

Method number two above would require the conversions of actual thrust required to corrected thrust at corrected RPM. Because you are starting with corrected thrust, which is equal to static gross thrust at static sea level, you only need to find net thrust.


For example, using data from the JT9D above.

Mach 0.0 100%cN1 RPM = 25,200 F net
Mach 0.1 100%cN1 RPM = 23,300 F net
Mach 0.2 100%cN1 RPM = 21,300 F net
~~~
Mach 0.7 100%cN1 RPM = 14,000 F net
Mach 0.8 100%cN1 RPM = 13,900 F net


So the chart for your engine will only vary from static gross thrust by the reduction for ram drag.


Because the corrected system accounts for ram rise during conversion, there is no need to account for that in the chart.


So to answer your question depends on if you have a chart that shows net thrust chart or not...

If not, do you have a static Mass Airflow chart?

If not, do you have a drag polar?

If not, let me know. There are a few ways to go about it but I can get long.

 

[Heretic]

If not, do you have a static Mass Airflow chart?

If not, do you have a drag polar?

No, nothing.

All I have is my line for mach 0 and these tables:
http://boeing-727.com/Data/engine/Cruise Tat -15.PDF
http://boeing-727.com/Data/engine/LRC-15 Chart.PDF
http://boeing-727.com/Data/engine/engdata.html

 

[jx_]

You will need to make a chart for the engine's massflow rate.

Or you will need cruise charts based on mach with fuel flow (...your PDFs)

-or a little of both-


This is not going to be an exact science because you don't have enough information. It will only be as accurate as the info you have to start with, but you should be able to get a good estimate. I don't know much about the JT8D, but I know it is a big family with many different engine models. This will make it harder to narrow things down.


Here's what you need to do to map the airflow.


You will need to find or make:

- exact area of the engine CORE exhaust, nozzle minus shaft. ( Ae )
- accurate static gross thrust chart at all CN1 points. ( Fg )
- accurate static fuel flow chart at all CN1 points ( Wf )


We are trying to solve for Mass Airflow/g (Wa/g)

***Note: Wa/g is Weight of massflow in pounds per second over gravitational velocity g in feet per second. g = 32.174. The value of Wa/g is currently less than 50. (The biggest jet engine in the world, the GE90, has a maximum massflow of 1500 lbs/s. 1500 / 32.174 = 47 Wa/g)

It is very important to get the exit area correct. This included the necessary deductions for the center shaft, casing, and struts. See the attached image for clarification.




Make a spreadsheet called Static Engine Parameters that looks like:

CN1____GrossThrust (Fg)____FuelFlow (Wf)____MassAirflow/g (Wa/g)
10 XXX
20
...
...


Fill in what you know, leave blank what you don't.


Here's the equation you'll need:

Step 1) Fg / sqrt( Fg / (Ae * 0.002377) ) = Ve = Exit Velocity

Step 2) Fg / Ve = Wa/g = Mass Airflow/g

So...

Fg / ( Fg / sqrt( Fg / (Ae * 0.002377) ) ) = Wa/g = Mass Airflow/g


Go through your table and fill in all the Wa/g lines using the equation.




We can crosscheck, or if you don't have the exit area available:

EPR * 2116 = Pt7

Pt7 - 2116 = q

sqrt( q / 0.5 / 0.002377 ) = Exit Velocity, Ve

Fg / Ve = Wa/g = Mass Airflow/g



***Note to other readers. These will only work on very low-bypass ratio engines like the JT8D. Modern high bypass engines have two velocities and two mass flows. In that case you need the velocity of the core exit (turbine exit velocity) and the velocity ratio or bypass velocity in order for the math to work.

Edit: There is a post added below that shows a chart made by Heretic. http://fsdeveloper.com/forum/showpost.php?p=656263&postcount=18 Notice how linear and scaled his chart is compared to the example thrust chart I've posted above. The difference is the mass and velocity of the bypass stream. The high bypass chart above has a much higher amount of ram drag, as well as a greater variation in the thrust curves at any given RPM.

Now, make a new table called Ram Drag by Mach

Mach____AircraftVelocity (V0)____Wa/g____Ram Drag (Fr)
0
0.1
0.2
0.3
....
...



Mach 1.0 Velocity at ISA sea level is 1116 ft p/s so

Mach * 1116 = V0 = Ambient flight velocity


Wa/g * V0 = Fr = Ram Drag


Complete that table.


Now, using Fg from the first table, you can make a new table call net thrust by mach.


CN1____Mach0.0____Mach0.1____Mach 0.2 ...
10 ____Fn
20
30
...


Subtract Fr at Mach from Fg at Cn1.

Fg - Fr = Fn




Once this is done, we need to cross check our work.


Pull up your cruise chart and pick an even mach that you have in your table. Find the fuel flow and N1 at that mach. Correct that fuel flow using:

Fuel Flow PPH (Wf) / (Delta * Theta2^0.5) = cFF = Corrected Fuel Flow in PPH

Correct the N1 using:

N1/Theta2^0.5 = cN1

Now compare your pairs of cFF/cN1 from this cruise chart against the pair listed in your first table called Static Engine Parameters. Is there an error of more than a hundred pounds or so?

If no major discrepancy is found, you can consider that data point reliable. Rerun the test at several mach and several CN1.



If there is an error found let me know and we'll go from there.









The only data I could find for the JT8D was takeoff exhaust velocity at 15587 lbs Fg = 1400 ft p/s Ve at 2.09 EPR and 360 Wa. Using this data I can give you a few quickie examples...

First, 360 Wa = 11.2 Wa/g


(2.09 EPR * 2116) - 2116 = 2306 q; sqrt( 2306 q / 0.5 / 0.002377 ) = 1393 Ve

1400 Ve ~~ 1393 Ve. This checks out.

So,

15587 Fg / (1393 Ve * 0.002377 * Area?) = 1393 Ve

15587 Fg / 1393 Ve = 11.2 Wa/g, so (1393 Ve * 0.002377 * Area?) should equal 11.2 Wa/g as well.


Thus,

15587 Fg / sqrt( 15587 Fg / (1393 Ve * 0.002377 * 3.3825 Ae) ) = 11.2 Wa/g

Ae = ~~3.3825

Wa/g = 11.2
Ve = 1393


At mach 0.8:

V0 = 661.478 * 0.8 = 812

Wa/g * V0 = Ram Drag

11.2 * 812 = 9094 Fr

So at Mach 0.8, Table 1506 thrust would be;

15587 - 9094 = 6492 Fn

This entry would be at the same CN1 as 15587 Fg.



Questions?

 

[jx_]

Googled a few things...

maybe these can help...

http://home.comcast.net/~flightsim/727_engines.htm

http://home.comcast.net/~flightsim/JT8D_engine_variants.htm



Good luck!

 

[Roy Holmes]

Heretic

The only disadvantage is that there's no FDE section (yet).)

There is a FDE Section in the wiki, look under aircraft design.

All it contains is a reference to my paper about 1502 to 1507 which is in the download section.
Roy

 

[jx_]

Hello Roy.

 

[Heretic]

Only got around doing this just now.
Got 326 lb/s for W (thank you, Flight Global*).


You should have noted that g is, as usual, the gravity constant, but this time in ft/s.
Since I only know g in SI-units, this had me thoroughly confused at first.



Still setting up the Fn table. Stand by...



- Edit:

Done with the table. My value for your example is in the same ballpark for a baseline of 16000 lbs flat at t/o ISA SL.



- Edit²:

This is my final table:



The formatting for the thrust factors facilitates copy+pasting into AAM readable files.


*http://www.flightglobal.com/pdfarchive/view/1978/1978 - 0044.html

 

[Roy Holmes]

http://www.jet-engine.net/civtfspec.html has a lot of JT-8D variants and each one has values for W. Different thrust would give different W. Values from 315 to 480

Roy

 

[jx_]

That looks great! Looks like you are getting the idea. Run some test flights (or spreadsheet simulations) and see if you can crosscheck the data.

One more hint, you can set the reference thrust in the aircraft config to 1.0 pound. Then table 1506 can be entered in lb/f directly, but the table converter you've made is excellent.


And yes, sorry about not explaining g. I will make an edit, thanks. I've always had a bad habit of that!