[Mustang]

HSFX 5.0 is oriented to SEOW !

http://wiki-seow-en.swil.fr/index.php/Main_Page


And have Aachens Flight Models - JGSME MOD

HSFX 5.0 Readme

Aachens Flight Models

A little Background:-

Aachen is a professional Aircraft design engineer, we were not sure if we wanted to go in this direction at first, but were so impressed by how much closer to what we have read flying some of these aircraft and fighting in them has come, that it was inconceivable to go back.

We hope that you share the views of the many 'test-pilots'. You retain a choice still anyhow. Using the full work is strictly optional.

Note:-
In Stock Configuration Oleg's Flyables are all Stock (by Oleg or TD) although all new aircraft * that Aachen has worked on have his FM's also.

In 'Expert Mode' All of the aircraft below use Aachens FM's.
We have altered the Netcode and the encryption for this mode, so that players cannot use the modified ‘Stock’ FM’s on stock servers.

Those that play online require very reasonably that everybody is playing with the same.
Expert mode is to protect this desire.
List of modified aircrafts

The complete list of planes modified is as follows:
Bf109:
E1, E1B, E3, E3b, E4, E4B, E4N, E7, E7N, E7NZ
F2, F2B, F2trop, F4, F4B, F4trop, F4Z
G1, G2, G2trop, G3, G4, G4trop, G5, G6early, G6Erla, G6Late, G6Mid, G6trop, G10, G10C3, G10Erla, G14, G14early, G14AS
K4, K4C3, K6, K14

Fw190:
A1, A2, A3, A4, A5, A5-165, A7, A7sturm, A8, A9
F8
D9, D9late, D11, D13

G55:
G55, G55late, G55-ss0, G55-ss0late

MC:
MC200, MC202, MC205, MC205V

P51:
P51B, P51C, P51CM, P51D20, P51D20NT, P51D25, P51D30

P40:
P40B, P40C, P40E, P40E-M-105, P40M

P38:
P38J, P38J10LO, P38J15LO, P38J25LO
P38L, P38L5LO, P38Llate

P47:
P47D10, P47D22, P47D27, P47D27late

Foreword
The modifications of flight and engine models presented in this work have started with an analytical evaluation of aircraft performances. In the following paragraphs a short description of the methodologies adopted in the analytical study can be found, specifically for the evaluation of aircraft polars.

Wing and tail polar
Are computed by adopting lift line theory (Weiselberger) using non linear section lift data (J.C. Sivells, R.H. Neely). Compressibility effects are taken into account. Normally, lift distribution, finite wing Cy and Cx computed are in very good agreement with computations performed according to DATCOM method (ref. E. Torenbeek, Synthesis of subsonic airplane design). This is due to the fact that studied aircraft configurations are un-swept and have high aspect ratios.


Figure 1 – Lift coefficient distribution on half wing span (Bf109G2 at SL 530km/h). Cyan line is result computed with iterative method (NACA Report 865) while yellow line is result computed with DATCOM method


Figure 2 – Lift coefficient distribution on half wing span (Bf109G2 at 1000m 250km/h 2g level turn). This condition illustrates the determination of stall-limited turn rate (in this case stall is incipient at 0.6 x half-wingspan). A tolerance of 0.05 g has been used to predict ultimate wing load factor for both stall-limited and power-limited turn rates.

Fuselage polar
Drag computation for fuselage has been performed by using slender body formulation (ref. E. Torenbeek, Synthesis of subsonic airplane design). Lift induced drag is accounted for in the computation. Formulation for fuselage lift induced drag is given in referenced document.

Propeller
Propeller performance computations have been performed by means of blade element theory. In the present document, since no detailed description of propeller blades was available, the blade section has been assumed to be a flat plate. Optimal propeller (i.e. blade twist) has been computed in the condition of 100% throttle at sea level. Hence the propeller has been analysed for all beta angles in the range specified in EMD (propPhiMax and propPhiMin) at maximum propeller revolutions (constant rpm propeller), thus obtaining propeller efficiency curve at full power rpms.
It should be noted that the assumption made on blade section leads to under-estimation of propeller efficiency (up to 5% at maximum speed) thus leading to a conservative estimation of aircraft performance.

Propeller slipstream
Is computed using blade element theory adopted for propeller performances estimation. It is worth mentioning that actuator disc theory produces very similar results in terms of slipstream velocity and mass flow rate. This is due to the fact that considered propellers have low loading factor. For the purpose of this study the complete fuselage, radiators (under-wing and under-fuselage), inner wing section and tail assembly are considered to be completely inside the propeller slipstream. The inner wing section area enveloped by propeller slipstream has been computed considering the propeller radius/wing span ratio. This assumption leads to a slight over estimation of wing drag since propeller slipstream tube has a contraction after the propeller (about ¼ - ½ of propeller radius downstream of propeller) to its final radius.
Small summary of modifications – Aircarft polars

dCl/dα has been evaluated according to the following formula:

Clα = f Clαth /(E+Clαth/(π AR)) [rad-1]

where Clαth is the 2D section lift coefficient derivative and E=1+(2 TR)/(AR (1+TR))

Drag coefficient second derivative has been evaluated according to the following formula:

d2Cd/dα2 = Clα2/(π AR e)

Second derivative of drag coefficient has been corrected with twist factor.

Clmax has been computed by computing Cl spanwise distribution and assuming linear spanwise variation of 2D section Clmax (ref. example figure below):



Bf109 slats
Bf109 slats has been treated as follows:
according to literature (R&M 2361 [sept. 1940]) slats open at Cl approximately 0,85-0,95. Second order Cd derivative for complete wing with slats deployed is computed at 5,3E-4. In the following figure the Cd as function of α is reported.

Since it is not possible to impose the Cd jump corresponding to slat open condition, the Cd is simulated with a second order derivative of 5,8E-4 with 0,8ｰ offset (ref. figure below).

This approximation limits the error in Cd estimation within +5% immediately before and -5% immediately after slat opening. Error tends to 0 moving away from slat openin threshold.
P51s CoG

In the models presented in this work, the P51 CoG position has been moved forward to replicate the position of the CoG in the configuration with 25 gallons in the 85 gallons fuselage fuel tank. From literature data the CoG for P51D configuration with 25 gallons in the 85 gallons fuselage fuel tank is 28.3% MAC. The P51s with full 85 gallons fuselage fuel tanks were statically unstable and the normal operating procedures for planes in such a configuration demanded to empty the 85 gallons fuselage fuel tank before all other tanks. At anything below 35 gallons, the P51s equipped with 85 gallons fuselage fuel tank were both statically and dynamically stable [America Hundred Thousands et al.]. Since the simulator does not allow for CoG movement with regards to fuel usage, and since the unstable configuration reproduced in the original models was deemed too conservative, it has been decided to adopt a statically and dynamically stable configuration as normally happened during combat operations. It is advisable to adopt a maximum fuel load of 75%.

P47D27 Late

In the models presented in this work, the P47D27Late has been modelled to reproduce (as best as technically possible) the flight characteristics and performances of P47M.