tail moment imbalance

[emir_]

L=CL×0.5×ρ×V^2×S
I want to design a V tail. The moment of the wing turned out to be -9.3 Nm. The tail needs to balance this moment. I'm thinking of taking the tail moment arm as 0.5-0.6m. I calculate a lift force with the formula I wrote, but my lift force turns out to be very low. Is the formula wrong or am I making a mistake somewhere else? My root chord length is 0.2m, my end chord length is 0.09m

 

[3DDave]

The way to do this is to pick a projected area of the stabilizer and then calculate the required Cl of the tail and from that determine what the angle of attack for the stabilizer should be to reach that Cl; note that the tail will operate in the downwash from the wing, so the actual angle of attack for the tail will be more negative than the angle of attack for the wing.

 

[emir_]

The way to do this is to pick a projected area of the stabilizer and then calculate the required Cl of the tail and from that determine what the angle of attack for the stabilizer should be to reach that Cl; note that the tail will operate in the downwash from the wing, so the actual angle of attack for the tail will be more negative than the angle of attack for the wing.

What are the ways to determine Cl?

 

[rb1957]

IDU ?

why the equation of lift ?

if you've calculated a Wm and you know the tail arm, then you know the tail force. How it manifests on a Vee tail is one matter. I'd say that the lift (of the inclined surface of the Vee tail) would have a vertical component equal to the required tail reaction. So if you had a 90 deg Vee tail, surfaces inclined at 45 deg, tail force required is 20N, lift on one Vee tail surface is 14.14N (which has a vertical component of 10, 2 surfaces = 20N).

But ... tail force adds to wing lift, which will add to wing moment. To clarify, nz*Wt = L - Lt, or L = nz*Wt+Lt (and of course there are many other terms, but I'm not trying to give a loads course).

 

[3DDave]

Cl=0.5×ρ×V^2×S/L is how the Cl of the tail is determined.

Then go to the chart of Cl vs. Alpha to see what Alpha is required to get that Cl. Note that since the tail surface is angled you need to deal with the true alpha for the airfoil section which is reduced by the amount of the V angle.

As rb1957 is hinting at, this begins the iterative loop of adding the downforce from the tail to the lift required of the wing, getting a new moment, which gives a new lift requirement for the tail.

The loop should converge pretty quickly.

You will also find that the tail lift (negative against the moment from the wing) also contributes moment and both form and induced drag.

 

[WKTaylor]

FAA AC 23-9 - [Large AC] Evaluation of Flight Loads on Small Airplanes with T, V, +, or Y Empennage Configurations

Design Process: V-Tails https://www.kitplanes.com/design-process-v-tails/

NASA-CR-195496~Aircraft Empennge Structural Detail Design

UWS 1 Design: V Tail Loads Part 1, Span Distribution

NASA-CR-195496~Aircraft Empennage Structural Detail Design

Aircraft Design for Freshmen Ch8

https://www.researchgate.net/profile/Aishwarya-Dhara/publication/339675700_Aircraft_Design_Guidelines_for_Freshmen/links/61b70efefd2cbd720098798e/Aircraft-Design-Guidelines-for-Freshmen.pdf

 

[rb1957]

Da-amn !!