PARAGLIDER DESIGN HANDBOOK

CHAPTER 4. LONGITUDINAL EQUILIBRIUM

4.1 Introduction and definitions

4.2 Equations of static equilibrium

4.2.1 Sum of moments from a point is equal to zero

4.2.2 Sum of vertical forces is zero

4.2.3 Sum of horizontal forces is zero

4.3. Solving the system

4.4. Spreadsheet

4.5. Simplified equilibrium aproximation

4.6. Empirical methods

1. Introduction and definitions

The
precise calculation of the
balance of the wing in flight is one of
the priorities of the designer of paragliders. This chapter details the
calculation of the balance from a two-dimensional analysis. It also
provides a simplified method and even empirical type.

The longitudinal study of 2D equilibrium in a paragliding, developed from the attached drawing. The method is based in references [2] and [3], using the same notations.

Fig. 1: Longitudinal equilibrium (click to enlarge).

Forces, points, angles, and lengths involved. Vector represent bold.

Forces:

P = lift

T = drag

M = momentum of aerodynamic forces

Ts= drag of lines

Tp= drag of pilot

Wp = weight of pilot

Wa = weight of wing

Points:

A = leading edge

B = trailing edge

Cp = center of aplication of the aerodynamic forces

P = pilot position

C = calage point

F = foyer aerodynamic

Ga = center of gravity of wing

Angles:

γ = angle of glide

θ = angle between horizon and airfoil chord

α = angle of atack (AoA)

Lengths:

L = length of lines

dist(A,B)= l

dist(A,Cp) = δ

dist(A,C) = σ

dist(Cp,C) = ε

Understanding
each of these elements
is important to understand the
balance of paragliding.

The paraglider is represented at a central profile representative of all the profiles of the wing. The geometry and aerodynamic properties of this profile represent the whole wing with sufficient accuracy, which is simplifying the model.

Forces P, T, amd momentum M, was the resultant of erodynamic forces acting on the profile, and applied in the center of pressure Cp. The values of aerodynamic coefficients Cz, Cx, Cm corresponding to different angles of attack α can be obtained through numerical models (program XFOIL).

Forces Ts and Tp depend on the lines scheme (diameters, lengths, pads) and the position of the pilot and his harness.

Wp and Wa apply in the centers of mass of the pilot and the wing to form a single center of mass, which should be balanced in flight under the cebter of pressure Cp of wing.

The calage point C, is an arbitrarily point defined by drawing a line perpendicular to the chord profile from the pilot position P. The AC distance is very important because it defines the relative lengths of lines and the behavior of the paraglider. Our main goal is to find the most appropriate calage.

The foyer point F. The foyer of not intervening directly in the calculations of balance, but their knowledge is important to know the stable or unstable profile chosen. Its definition is not simple, and can be performed as follows: The point of the chord from the moment produced by the coeficient of lift (Cz) balances (equal and opposite sign) the aerodynamic moment Cm. That is, the distance between the foyer F and the center of pressure Cp, multiplied by the coefficient of lift Cz is equal to Cm. For practical purposes it is enough to know that the foyer is located at 25% of the chord from LE in most profiles.

The angle of
glide γ is the angle between the horizontal line and
flight path. The paraglider is represented at a central profile representative of all the profiles of the wing. The geometry and aerodynamic properties of this profile represent the whole wing with sufficient accuracy, which is simplifying the model.

Forces P, T, amd momentum M, was the resultant of erodynamic forces acting on the profile, and applied in the center of pressure Cp. The values of aerodynamic coefficients Cz, Cx, Cm corresponding to different angles of attack α can be obtained through numerical models (program XFOIL).

Forces Ts and Tp depend on the lines scheme (diameters, lengths, pads) and the position of the pilot and his harness.

Wp and Wa apply in the centers of mass of the pilot and the wing to form a single center of mass, which should be balanced in flight under the cebter of pressure Cp of wing.

The calage point C, is an arbitrarily point defined by drawing a line perpendicular to the chord profile from the pilot position P. The AC distance is very important because it defines the relative lengths of lines and the behavior of the paraglider. Our main goal is to find the most appropriate calage.

The foyer point F. The foyer of not intervening directly in the calculations of balance, but their knowledge is important to know the stable or unstable profile chosen. Its definition is not simple, and can be performed as follows: The point of the chord from the moment produced by the coeficient of lift (Cz) balances (equal and opposite sign) the aerodynamic moment Cm. That is, the distance between the foyer F and the center of pressure Cp, multiplied by the coefficient of lift Cz is equal to Cm. For practical purposes it is enough to know that the foyer is located at 25% of the chord from LE in most profiles.

The angle θ is angle between the horizon and the airfoil chord.

The angle α is the angle of atack (AoA) formed between the chord and the trajectory of the wing.

2. Equations of static equilibrium

There are three classical equations of equilibrium in two dimensions:

Σ M = 0 The sum of moments from a point is equal to zero.

Σ V = 0 The sum of vertical forces is zero.

Σ H = 0
The sum of horizontal forces
is zero.

2.1
Sum of moments
from a point
is equal to zero.

Σ M = 0

By convention we will take moments with respect to the leading edge point A

M_{A}(P) + M_{A}(T) + M_{A}(Ts) + M_{A}(Tp) + M_{A}(Wp) + M_{A}(Wa) + M = 0

(1) + (2) + (3) + (4) + (5) + (6) + (7) = 0

Moments positive counter-clockwise. We
analyze in detail each
term:

(1) = q_{0}
* Cz * δ * cos (α)

q_{0
= (1/2) * }ρ * S * V^{2 }was
the dynamic pressure, where ρ is the density of air, S the area
of the wing, and V the velocity.

P = k * Cz where Cz is the coeficient of
lift and k was unitary vector
ortogonal to trajectory

(2) = q_{0}
* Cxa * δ * sin (α)

T = i * Cxa where Cxa is the coeficient of drag of the wing and i was unitary vector along trajectory

Cxa = Cxoa + Cxi

where Cxoa is the drag of shape and friction of wing

and Cxi is the indiced drag Cxi=Cz^{2}(α)/(π
* λ * e)

λ = b^{2} / S is the aspect ratio, b span and
S surface

e ≈ 0.9 Oswald factor

(3) = (L/3) * Ts * cos (α) + σ * Ts * sin
(α) = q_{0 }*_{ }C_{xs
}* ((L/3) * cos (α) + σ * sin
(α) )

where Ts = q_{0 }*_{ }Cxs
is the drag of lines

(4) = q_{0 }*_{ }Cxp
* (L * cos (α) + σ * sin (α) )

where Tp = q_{0 }*_{ }Cxp
is the drag of the
pilot-harness

(5) = - m_{p}
* g * ( L * sin (θ) + σ * cos (θ) )

Wp = m_{p}
* g is wheigt of the pilot, g=9.81 m/s^{2
}

(6) = - (l/3) * m_{a}
* g * cos (θ)

Wa = m_{a}
* g and the center of gravity of the wing in
the third the chord (l/3)

(7) = - q_{0 }*_{ }Cm
is the aerodynamic moment

Making substitutions in each of the terms, the equation of equilibrium of moments becomes:

q_{0} * Cz * δ * cos (α) +

q_{0} * Cxa * δ * sin (α) +

q_{0
}*_{ }C_{xs }*
((L/3) * cos (α) + σ *
sin (α) ) +

q_{0
}*_{ }Cxp
* (L * cos (α) + σ * sin (α) ) +

- m_{p} * g * ( L * sin (θ) + σ * cos (θ) ) +

- (l/3) * m_{a}
* g * cos (θ) +

- q_{0
}*_{ }Cm
= 0

2.2
Sum of vertical
forces is zero.

Positive in the direction of gravity.

Σ V = 0

m_{p} * g + m_{a}
* g +

- q_{0} * Cz * cos (γ) - q_{0}
* Cxa * sin (γ) +

- q_{0} * Cxs * sin (γ) - q_{0}
* Cxp * sin (γ) =
0

then,

g * (m_{p }+
m_{a} ) - q_{0}
* (
Cz * cos (γ) + ( Cxa + Cxs + Cxp) * sin (γ) ) = 0

Cxa + Cxs + Cxp = CxT is the total drag coeficient

g * ( m_{p }+
m_{a} ) - q_{0}
* ( Cz *
cos (γ) + ( CxT) * sin (γ) ) = 0

2.3
Sum of horizontal
forces is zero.

Positive left to right

Σ H = 0

q_{0} * Cxa * cos (γ) + q_{0}
* Cxs * cos
(γ) + q_{0} * Cxp * cos (γ) - q_{0}
* Cz * sin (γ) = 0

then,

CxT * cos (γ) = Cz * sin (γ)

or,

tan (γ) = CxT / Cz

Glide ratio ( finesse ) = 1 / tan (γ) by definition

then GR = Cz / CxT

Graphical notes:

Fig. 2: Note 1

Fig. 3: Note 2

3. Solving the system

Equilibrium equations:

[1]

q

q

q

q

- m

- (l/3) * m

- q

[2] g * ( m

[3] tan (γ) = CxT / Cz

Auxiliar equations:

[4] q

[5] Cxa = Cxoa + Cxi

[6] Cxi=Cz

[7] λ = b

[8] CxT = Cxa + Cxs + Cxp

[9] Cxa = Cxoa + Cxi

[10] γ = α + θ

[11] GR = Cz / CxT

Analytical solution (to be developed)

(...)

Numerical method:

The solution of the system can be obtained by analytical or by numerical method. We use the numerical iterative way, based on the balance equation of moments to get their sum equal to zero. The objective is to obtain a combination of parameters to meet the equations and provide the desired angle of glide.

The solution of the system can be obtained by analytical or by numerical method. We use the numerical iterative way, based on the balance equation of moments to get their sum equal to zero. The objective is to obtain a combination of parameters to meet the equations and provide the desired angle of glide.

Of the variables included in the equations above consider the following:

Inputs

m

numerical estimates

Cxs, Cxp

aerodynamic values obtained numerically

Cz, Cx, Cm, Cp

Results

V, γ, δ, θ, GR, Cxi, Cxa, CxT, q

4. Spreadsheet

SOLVING THE EQUILIBRIUM
EQUATIONS |
||||||||

Wing name: | LAB_exa | |||||||

Airfoil | Laboratori | |||||||

Surface | 25.70 | m2 | ||||||

Span | 10.25 | m | ||||||

Chord | 3.00 | m | ||||||

Considered chord | 3.00 | m | ||||||

Aspect ratio | 4.09 | |||||||

Lines length | 6.00 | m | ||||||

Pilot masse | 80.00 | Kg | ||||||

Wing masse | 5.00 | Kg | ||||||

G | 9.81 | m/s2 | ||||||

RHO | 1.11 | Air density | ||||||

Oswald | 0.90 | Oswald coeficient | ||||||

CX pilot | 0.015 | Pilot drag | ||||||

CX lines | 0.020 | Lines drag | ||||||

Alfa | 5.00 | Deg | 0.087 | Rad | Flight angle | |||

Cz(alfa) | 0.900 | Lift | ||||||

Correcció 3D Cz | 0.624 | |||||||

Cz(alfa) correction 3D | 0.562 | |||||||

Cx0a(alfa) | 0.009 | |||||||

Cm(alfa) | -0.042 | |||||||

Cm0 | -0.038 | |||||||

Alfa0 | -1.700 | |||||||

Cxi(alfa) | 0.070 | cxo+ci | Cxo | Ci | Cp | Cs | ||

CxT | 0.114 | Total drag | 100.00% | 69.24% | 7.65% | 61.59% | 13.18% | 17.58% |

GAMMA(alfa) | 0.200 | Rad | 11.45 | Deg | 4.94 | GR | (Cz3D) | |

8.57 | GR max | |||||||

Theta | 0.113 | Rad | 6.45 | Deg | ||||

Q0(alfa) | 1455.1 | |||||||

V | 10.09 | m/s | 36.33 | km/h | Speed | |||

Bl | -0.046 | Cm/Cz | ||||||

Xcp | 0.296 | Cp (%1 chord) | ||||||

CP | 0.889 | Cp position (m) | ||||||

Calage | 0.433 | Calage (m) | ||||||

Sigma | 0.144 | Calage chord fraction | ||||||

Euilibrium of moments to LE | ||||||||

M1= | -725.02 | Moment RFA | ||||||

M2= | -59.08 | Moment lines | ||||||

M3= | -131.29 | Moment pilot drag | ||||||

M4= | 866.82 | Moment pilot weight | ||||||

M5= | 48.74 | Moment wing weight | ||||||

SUM= | 0.17 | Zero in equilibrium | ||||||

FH= | 0 |
Zero in equilibrium | ||||||

FV= | 0 |
Zero in equilibrium |

Note: The spreadsheet is still in development (there are some details to be corrected). When functional the file was posted in .ods .xls format for download.

5. Simplified equilibrium aproximation method

A
more simplified approach, although
based on the same principle of equilibrium is described below:

1. We draw the central wing profile in the planned equilibred flight.

2. We assume that the wing is flying with a glide ratio of GR (as expected for our wing), which corresponds to an angle γ

3. We assume
that the airfoil incidence corresponding to max
glide ratio or slightly higher (10%), is α. Depending on the profile of
the whole wing, usually those angles ranging between 5 deg and 10 deg.
It is preferable to a plan to estimate the glide at the highest speed,
so we start by estimating the angle at 5 deg.1. We draw the central wing profile in the planned equilibred flight.

2. We assume that the wing is flying with a glide ratio of GR (as expected for our wing), which corresponds to an angle γ

GR | 1 | 2 | 3 | 4 | 5 | 6 | 7 | 8 | 9 | 10 | 11 | 12 |

Angle γ (deg) | 45 | 26.57 | 18.43 | 14.04 | 11.31 | 9.46 | 8.13 | 7.13 | 6.34 | 5.71 | 5.19 | 4.76 |

4. Then draw the representative profile of the wing in flight. The angle of the profile chord and the horizon is θ = γ - α

5. We locate the position of center of pressure Cp for the chosen profile and angle of incidence α.

6. We locate the position of the pilot P in the same vertical of the center of pressure Cp, at a distance equal to the lines length.

7. From the pilote point P draw a perpendicular to the line of chord wing to determine the calage C of the wing.

Fig. 4: The simplified equilibrium method

The method is simple and can be used for a first approximation. But has the disadvantage of having to estimate a priori the glide angle (deductible by the general characteristics of the wing) and the angle of attak at maximum glide ratio (more difficult to estimate). Some corrections on the length of the riser or lines in the prototype are necessary to determine the calage just ideal.

6. Empirical methods

Many designers use empirical methods based on their experiences with previous models. These methods are simple and effective.

A very effective method is to study the profile and the relative
length of the lines of a existing paraglider with similar
characteristics to
the project. We only need to study the middle section and deduct
graphically the
calage employee, with and
without speed system.