Coefficient of Pitching Moment Respect to Fuselage Contribution Calculator

✖The Wing Area is the projected area of the planform and is bounded by the leading and trailing edges and the wing tips.ⓘ Wing Area [S_w]			+10% -10%
✖The Wingspan (or just span) of a bird or an airplane is the distance from one wingtip to the other wingtip.ⓘ Wingspan [b]			+10% -10%
✖The Average Width of Fuselage refers to the typical diameter or width of the central body structure of an aircraft.ⓘ Average Width of Fuselage [w_f]			+10% -10%
✖The wing zero lift angle, relative to the fuselage reference line, refers to the angle formed between the chord line of the wing and the reference line of the fuselage.ⓘ Wing Zero Lift Angle [α_0,w]			+10% -10%
✖The Incidence of Fuselage Camber Line relative to the fuselage reference line refers to the angle formed between the camber line of the fuselage and the reference line of the fuselage.ⓘ Incidence of Fuselage Camber Line [i_f]			+10% -10%
✖The Length of Fuselage Increment refers to the increase in the longitudinal dimension of an aircraft's fuselage.ⓘ Length of Fuselage Increment [Δx]			+10% -10%

✖The Moment Coefficient for Fuselage Contribution is the sum of the contributions of each fuselage component, including the nose, cabin, and tail cone.ⓘ Coefficient of Pitching Moment Respect to Fuselage Contribution [Cm_0,f]

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Coefficient of Pitching Moment Respect to Fuselage Contribution Solution

STEP 0: Pre-Calculation Summary

Formula Used

Moment Coefficient for Fuselage Contribution = (1/(36.5*Wing Area*Wingspan))*sum(x,0,Wingspan/2,Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line)*Length of Fuselage Increment)
Cm_0,f = (1/(36.5*S_w*b))*sum(x,0,b/2,w_f^2*(α_0,w+i_f)*Δx)
This formula uses 1 Functions, 7 Variables

Functions Used

sum - Summation or sigma (∑) notation is a method used to write out a long sum in a concise way., sum(i, from, to, expr)

Variables Used

Moment Coefficient for Fuselage Contribution - The Moment Coefficient for Fuselage Contribution is the sum of the contributions of each fuselage component, including the nose, cabin, and tail cone.
Wing Area - (Measured in Square Meter) - The Wing Area is the projected area of the planform and is bounded by the leading and trailing edges and the wing tips.
Wingspan - (Measured in Meter) - The Wingspan (or just span) of a bird or an airplane is the distance from one wingtip to the other wingtip.
Average Width of Fuselage - (Measured in Meter) - The Average Width of Fuselage refers to the typical diameter or width of the central body structure of an aircraft.
Wing Zero Lift Angle - (Measured in Radian) - The wing zero lift angle, relative to the fuselage reference line, refers to the angle formed between the chord line of the wing and the reference line of the fuselage.
Incidence of Fuselage Camber Line - The Incidence of Fuselage Camber Line relative to the fuselage reference line refers to the angle formed between the camber line of the fuselage and the reference line of the fuselage.
Length of Fuselage Increment - (Measured in Meter) - The Length of Fuselage Increment refers to the increase in the longitudinal dimension of an aircraft's fuselage.

STEP 1: Convert Input(s) to Base Unit

Wing Area: 184 Square Meter --> 184 Square Meter No Conversion Required
Wingspan: 20 Meter --> 20 Meter No Conversion Required
Average Width of Fuselage: 3.45 Meter --> 3.45 Meter No Conversion Required
Wing Zero Lift Angle: 0.31 Radian --> 0.31 Radian No Conversion Required
Incidence of Fuselage Camber Line: 3.62 --> No Conversion Required
Length of Fuselage Increment: 5.12 Meter --> 5.12 Meter No Conversion Required

STEP 2: Evaluate Formula

Substituting Input Values in Formula

Cm_0,f = (1/(36.5*S_w*b))*sum(x,0,b/2,w_f^2*(α_0,w+i_f)*Δx) --> (1/(36.5*184*20))*sum(x,0,20/2,3.45^2*(0.31+3.62)*5.12)

Evaluating ... ...

Cm_0,f = 0.0196133917808219

STEP 3: Convert Result to Output's Unit

0.0196133917808219 --> No Conversion Required

FINAL ANSWER

0.0196133917808219 ≈ 0.019613 <-- Moment Coefficient for Fuselage Contribution

(Calculation completed in 00.020 seconds)

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< Fuselage Contribution Calculators

Pitching Moment Coefficient for Fuselage Contribution

LaTeX Go Moment Coefficient for Fuselage Contribution = (Final Correction Factor for Body Fitness Ratio-Initial Correction Factor for Body Fitness Ratio)/(36.5*Wing Area*Mean Aerodynamic Chord)*int(Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line),x,0,Incidence of Fuselage Camber Line)

Coefficient of Pitching Moment Respect to Fuselage Contribution

LaTeX Go Moment Coefficient for Fuselage Contribution = (1/(36.5*Wing Area*Wingspan))*sum(x,0,Wingspan/2,Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line)*Length of Fuselage Increment)

Coefficient of Pitching Moment Respect to Fuselage Contribution Formula

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What is Coefficient of Pitching Moment Respective to Fuselage Contribution ?

The Coefficient of Pitching Moment Respect to Fuselage Contribution, often denoted as represents the derivative of the pitching moment coefficient with respect to changes in the angle of attack induced by the fuselage alone. This coefficient quantifies the contribution of the fuselage to the overall pitching moment experienced by the aircraft.

What is Positive and Negative Pitching Moment refers to?

A positive value indicates that the fuselage contributes to a nose-up pitching moment, while a negative value indicates a nose-down contribution. This coefficient is essential for understanding the overall aerodynamic behavior and stability of the aircraft, particularly in relation to the fuselage's impact on pitching moments during various flight conditions.

How to Calculate Coefficient of Pitching Moment Respect to Fuselage Contribution?

Coefficient of Pitching Moment Respect to Fuselage Contribution calculator uses Moment Coefficient for Fuselage Contribution = (1/(36.5*Wing Area*Wingspan))*sum(x,0,Wingspan/2,Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line)*Length of Fuselage Increment) to calculate the Moment Coefficient for Fuselage Contribution, The Coefficient of pitching moment respect to fuselage contribution, often denoted as fuselage measures the change in pitching moment per unit change in a parameter related to the fuselage. This parameter could be the angle of attack, sideslip angle, or another relevant factor. Moment Coefficient for Fuselage Contribution is denoted by Cm_0,f symbol.

How to calculate Coefficient of Pitching Moment Respect to Fuselage Contribution using this online calculator? To use this online calculator for Coefficient of Pitching Moment Respect to Fuselage Contribution, enter Wing Area (S_w), Wingspan (b), Average Width of Fuselage (w_f), Wing Zero Lift Angle (α_0,w), Incidence of Fuselage Camber Line (i_f) & Length of Fuselage Increment (Δx) and hit the calculate button. Here is how the Coefficient of Pitching Moment Respect to Fuselage Contribution calculation can be explained with given input values -> 0.019563 = (1/(36.5*184*20))*sum(x,0,20/2,3.45^2*(0.31+3.62)*5.12).

FAQ

What is Coefficient of Pitching Moment Respect to Fuselage Contribution?

The Coefficient of pitching moment respect to fuselage contribution, often denoted as fuselage measures the change in pitching moment per unit change in a parameter related to the fuselage. This parameter could be the angle of attack, sideslip angle, or another relevant factor and is represented as Cm_0,f = (1/(36.5*S_w*b))*sum(x,0,b/2,w_f^2*(α_0,w+i_f)*Δx) or Moment Coefficient for Fuselage Contribution = (1/(36.5*Wing Area*Wingspan))*sum(x,0,Wingspan/2,Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line)*Length of Fuselage Increment). The Wing Area is the projected area of the planform and is bounded by the leading and trailing edges and the wing tips, The Wingspan (or just span) of a bird or an airplane is the distance from one wingtip to the other wingtip, The Average Width of Fuselage refers to the typical diameter or width of the central body structure of an aircraft, The wing zero lift angle, relative to the fuselage reference line, refers to the angle formed between the chord line of the wing and the reference line of the fuselage, The Incidence of Fuselage Camber Line relative to the fuselage reference line refers to the angle formed between the camber line of the fuselage and the reference line of the fuselage & The Length of Fuselage Increment refers to the increase in the longitudinal dimension of an aircraft's fuselage.

How to calculate Coefficient of Pitching Moment Respect to Fuselage Contribution?

The Coefficient of pitching moment respect to fuselage contribution, often denoted as fuselage measures the change in pitching moment per unit change in a parameter related to the fuselage. This parameter could be the angle of attack, sideslip angle, or another relevant factor is calculated using Moment Coefficient for Fuselage Contribution = (1/(36.5*Wing Area*Wingspan))*sum(x,0,Wingspan/2,Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line)*Length of Fuselage Increment). To calculate Coefficient of Pitching Moment Respect to Fuselage Contribution, you need Wing Area (S_w), Wingspan (b), Average Width of Fuselage (w_f), Wing Zero Lift Angle (α_0,w), Incidence of Fuselage Camber Line (i_f) & Length of Fuselage Increment (Δx). With our tool, you need to enter the respective value for Wing Area, Wingspan, Average Width of Fuselage, Wing Zero Lift Angle, Incidence of Fuselage Camber Line & Length of Fuselage Increment and hit the calculate button. You can also select the units (if any) for Input(s) and the Output as well.

How many ways are there to calculate Moment Coefficient for Fuselage Contribution?

In this formula, Moment Coefficient for Fuselage Contribution uses Wing Area, Wingspan, Average Width of Fuselage, Wing Zero Lift Angle, Incidence of Fuselage Camber Line & Length of Fuselage Increment. We can use 1 other way(s) to calculate the same, which is/are as follows -

Moment Coefficient for Fuselage Contribution = (Final Correction Factor for Body Fitness Ratio-Initial Correction Factor for Body Fitness Ratio)/(36.5*Wing Area*Mean Aerodynamic Chord)*int(Average Width of Fuselage^2*(Wing Zero Lift Angle+Incidence of Fuselage Camber Line),x,0,Incidence of Fuselage Camber Line)

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