karlbamforth,
What kind of wing beam do you have?
You could have a single closed wing box with front and rear spars.
You could also have a wing box made up by most (or part) of the wing with several webs forming a multiple closed cells beam.
In the case you have distinct spars, these could be made up of upper & lower flanges and webs (and vertical struts for stabilizing diagonal tension due to shear buckling), but you could also have a truss-structure instead of webs.
All these different possibilities require their specific approach. It is not so difficult, but you have to deal with several things like for example compression buckling of the flanges, combined compression & bending & shear buckling of the webs and all that goes with it like diagonal tension loads in the flanges and the struts, inter rivet buckling (if you have rivet connections of course) , also maybe you have to check fatigue loads etc... I don’t want to discourage you, as a matter of fact I would be happy to help you, but I have too much other stuff to do and this website is not meant for finding people but to share knowledge. I think it would be worthwhile to find a friend structural engineer and show him your problem.
You also need to define some load cases to load your wing beam with. I think you can get inspiration in subchapter C of the airworthiness regulations on what limit load cases you might consider. Once you have defined your most critical load cases, you should know per case what the translation and rotation accelerations are and the airspeeds. From these you determine the air loads and the inertial loads on your wings. If your wing is slender enough I would determine the inertial loads along the wing sections cg-line and the air loads along the wing section aerodynamic center (quarter chord) line.
A simple example.
A. Air Loads:
For the air loads, the lift forces will be most important.
To know this lift force you can proceed in several ways. For example if you know only the vertical aircraft acceleration and vertical gravitational acceleration component you can estimate the Lift force as follows.
To stay in a simple situation, for example for a vertical up gust case or a pull-up maneuver with wings level and fuselage horizontal,
the Lift on the wings = (Aircraft Mass)*(Aircraft vertical acceleration + g ).
By the way g is the gravitational acceleration (9,81 m/s2).
Now this lift is produced by the wing and has to be distributed according to a realistic wing lift distribution.
Lift = CL 0.5 rho V2 S,
Rho = air density (1,225 kg/m3 at sea level for standard atmosphere)
S = wing projected area (m2)
V is air speed (m/s)
and CL = integrated lift coefficient cl along the wingspan:
CL = 2* Integral from 0 to half span (cl

c

dy) / S
S = wing projected area, y = span coordinate, c

is local wing section chord, cl

is local wing section lift coefficient, dy is wing span increment.
The shape of cl

distribution is determined by your wing plan form = wing taper, wing twist and wing sweep. You can either look it up in a book like ‘Theory of wing sections’ By IRA H. ABBOOTT & E. VON DOENHOFF or calculate it with a panel program or just use common sense and take a approximate approach. For example suppose you have an elliptical lift distribution, then cl

= cl is constant. In that case CL = cl.
So cl

= cl = Lift / (0.5 rho V2 S),
thus the local lift load increment at span position y is dl

= cl * c

* dy * 0.5 rho V2
Besides the Lift distribution, you also have to find the Torque distribution on the wing, this is simply cm

= cmac at the wing section aerodynamic center (quarter chord) line .
thus the local torque load increment at span position y is dm

= cmac * c

*c

* dy * 0.5 rho V2
Concluding you have dl

forces upwards and dm

pitch down moments acting on the wing at span wise positions y and at the quarter chord line section

locations.
B. Inertial load:
In this simple situation of just a vertical acceleration, your distributed inertial loads along the wing span y are simply
(Section mass per span length at position y)*(Aircraft vertical acceleration + g )*dy.
This inertial load increment (inertail relief) is pointing downwards and acts at the mass cg’s of your wing section

at wing span position y.
If you want to be conservative and your wing does not contain any important masses, you can neglect this force. That is to say, if you can prove that your wing is strong enough when taking just the air loads without inertial relief, then the authorities will be even happier!
Good Luck,
OneMoreChance
PS soon I may also need help in the future in a question