SELECTION OF CRITICAL LOADS

METHOD

The  computer  program  SELECT.BAS reads the  data  for  all 
points on the V-n diagrams calculated from the data file  created 
by FLTLOADS.BAS. It searches for the critical flight loads on the 
major  components of the airplane structure, including the  wing, 
horizontal tail, vertical tail and fuselage. Critical aileron and 
flap loads are calculated later with separate programs.

INPUT

Much  detail concerning the geometry and inertia is  brought 
forward from the previous programs as input to this program.  The 
areo  surfaces and inertia influence every part of  the  airplane 
loads.

The  tail  surface loads are due to control  deflections  to 
react the wing and inertia loads or to produce pitching or yawing 
accelerations or attitude angles. The tail loads are more  simply 
evaluated aerodynamically than the wing since the regulations are 
conservative  in assuming that the spanwise distribution is  pro- 
portional to the chord, ignoring the fact that it tends toward an 
elliptical distribution. Also the airfoil does not influence  the 
slope of the lift curve nearly so much as the aspect ratio.  Thus 
the  slope of the lift curve tail surfaces may be assumed  to  be 
simply  a = 6.28/[1+6.28/(3.1416xAR)]. For the vertical  tail  of 
the sample airplane the aspect ratio is 1.52 making the slope  of 
the  lift curve 2.713 Cr per radian. The remaining input data  is 
available  from previously generated data or from the three  view 
drawing.

WING LOADS 
Up Loads PHAA, PMAA and PLAA

The  V-n data is searched for the largest resultant load  on 
the  wing for positive maneuver load factor at V> (PHAA),  at  V/-
(PMAA)  and  at Vp (PLAA) as required by  FAR  23.333(b) (1) .  The 
largest positive gust load at VQ required by FAR  23.333(c)(1)(i) 
is included in the PMAA search.

Down Loads NMAA

The V-n data is searched for the largest load on the wing 
for negative maneuver load factor at V^ (NMAA) required by FAR 
23.337 (b) (2). The largest negative gust at V^ required by FAR 
23.333(c)(l)(i) is included in the NMAA search. WARNING: This 
routine is not presently in the computer program. A visual check

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FAR 23 LOADS

of  the  V-n data should be made to verify that  actual  negative 
load is about .4 times the positive load and that the wing  nega- 
tive  bending strength is at least .5 times the strength  of .the 
wing  in  positive bending. For a strut based wing (such  as  the 
Piper  Cub or Cessna 172) the strut should be  substantiated  for 
the compressive column load in this condition.

Accelerated Roll

The  V-n  data is searched for the condition  producing  the 
largest  airload  at speed V^ to be  modified  for  unsymmetrical 
rolling  acceleration conditions of FAR 23.349(a)(l) or (2).  See 
page 103 for an explanation and example. See point 20 on page 186 
for a typical accelerated roll location on the V-n diagram.  This 
condition  is critical for shear across the wing  center  section 
and for attachment of heavy items in the wing.

Steady Roll

The  V-n data searches the steady roll conditions at  speeds 
V^, VQ and Vp for the maximum wing torsion condition required  by 
FAR 23.349(b) using the method of CAM 3.222(c) and (d). See pages 
103 and 104 for an explanation and example. The lift  coefficient 
distribution  may  be  assumed the same as  for  the  symmetrical 
flight  condition  at  C^=2/3 times the  positive  maneuver  load 
factor. The section pitching moment coefficient is calculated  as 
the the section pitching moment coefficient of the basic  airfoil 
plus an increment of .01 times the aileron deflection in degrees. 
This method is approved by FAR Policy statement CAM 3.191-l(a).

HORIZONTAL TAIL LOADS

Balancing Tail Loads

The  V-n  data  is searched for  the  largest  positive  and 
negative balancing horizontal tail loads with flaps retracted and 
with  flaps  extended as required by FAR 23.421(a) and  (b).  The 
distribution in figure B6 of appendix B may be used.

Unchecked Pull Up Maneuver

The  largest  unchecked pull up maneuver  (down  tail  load, 
elevator trailing edge up) at V^ is calculated as required by FAR 
23.423(a)(l)  using the average loading of B23.11 of Appendix  B, 
curves  (A, B and C) of Figure Bl and distribution of Figure  B7. 
(The  equations  producing the curves of Figure Bl  are  actually 
used in the computer program instead of reading Figure Bl.)  This 
load  occurs  at several cg's and altitudes since the  Figure  Bl 
does not consider altitude.

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SELECTION OF CRITICAL LOADS

However, a search is made for the balanced condition at 1  G 
that results in the largest unbalanced pitching moment in case  a 
refined analysis (including linear and angular inertia forces) of 
the aft fuselage is desired.

An FAA policy statement CAM 3.216-2(b) allows the down  load 
to  be carried forward to the wing attach points,  assuming  that 
the fuselage load factor equals zero. The moment at the wing need 
not be balanced out as a couple at the wing attach points.

Unchecked Push Down Maneuver

The largest unchecked push down maneuver tail load (up  tail 
load,  elevator  trailing edge down) at V^ is calculated  as  re- 
quired by FAR 23.423(a)(2) using the average loading of B23.11 of 
Appendix  B, curve B of Figure Bl and distribution of Figure  B7. 
To simplify the analysis, FAA Policy statement CAM 3.216-3 the up 
load on the horizontal tail may be carried through the attachment 
of  the horizontal tail surfaces to the fuselage and local  fuse- 
lage  members. No other structure need be investigated  for  this 
condition.

Checked Pull Up Maneuver

At  speeds  above Vx the down tail load  (elevator  trailing 
edge  up) is checked with a up tail load (elevator trailing  edge 
down) so that the maneuver load factor is not exceeded. The first 
part of the checked maneuver condition required by FAR  23.423(b) 
results in a normal acceleration of 1.0 and a pitching  accelera- 
tion  =  39N(N-1.5)/V where N = positive limit  maneuvering  load 
factor and V = initial speed in knots.

The  unbalanced tail load increment is calculated using  the 
equation T = la where T=increment times distance cg to tail, I  = 
airplane  pitching  inertia and a =  pitching  acceleration.  The 
total  tail  load is the sum of the increment and  the  balancing 
load at 1 G. The program searches for the largest down tail load.

The  maneuvering load increment in figure B2 of  Appendix  B 
and  the  distribution in figure B7 may be used according to  FAR 
23.423(B)  but  the more rational method above is  used  in  this 
program.

Checked Push Down Maneuver

The  the  up tail load (elevator trailing edge down) in  the 
push  down maneuver to check the pull up maneuver above  is  also 
required by FAR 23.423(b). This results in a normal load factor = 
positive  maneuver  load factor and a pitching acceleration  = 
39N(N-1.5)/V. The total tail load is the sum of the increment and

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FAR 23 LOADS

the  balancing  load at positive limit load factor.  The program 
searches for the largest up tail load.

The  maneuvering load increment in figure B2 of  Appendix  B 
and  the  distribution in figure B8 may be used according to  FAR 
23.423(B)  but  the more rational method above is  used  in  this 
program.

Up and Down Gust Tail Loads Flaps Retracted

The   horizontal  tail loads are  calculated  for  the  gust 
velocities  specified  in FAR 23.333(c) with flaps  retracted  as 
required by FAR 23.425(a)(1).

This  computer program does not use the average loadings  in 
figures  B3 and B4 of Appendix B which may be used  according  to 
FAR  23.425(b).  Instead  the initial balancing  tail  loads  for 
steady  unaccelerated  flight are added to the  incremental  tail 
load  calculated  by the rational method of  FAR  23.425(d).  The 
downwash factor de/da is assumed to be 36a^/AR^ per CAM  3.217(c) 
where a^ = slope of lift curve of wing (C^ per degree) and AR^  = 
aspect ratio of wing. Then the largest up and down tail loads are 
selected  as the critical up and down gust tail loads with  flaps 
retracted.

The following FAA policy applies to these conditions per CAM 
3.217-1. "The specified up gust and down gust load may be carried 
through  the  fuselage structure to the wing  attachment  points, 
assuming that the fuselage load factor is equal to that given  by 
positive  and  negative gusts of 30 fps at V/-  respectively.  The 
angular inertia forces in general produce relieving loads and may 
be taken into account if desired. The attachments of concentrated 
mass items in the rear portion of the fuselage may be  critically 
loaded by pitching acceleration forces."  The 30 fps and the gust 
formulas have changed since CAM 3, but the policy should not have 
changed.

Up and Down Gust Tail Load Flaps Extended

The horizontal tail loads are calculated for the gust veloc- 
ities  of 25 f.p.s. as specified in FAR 23.425(a)(2)  with  flaps 
extended   at   Vp  for  flight  conditions  specified   in   FAR 
23.345(a)(2).

This  computer program does not use the average loadings  in 
figures  B3 and B4 of Appendix B which may be used  according  to 
FAR  23.425(b).  Instead  the initial balancing  tail  loads  for 
steady  unaccelerated  flight are added to the  incremental  tail 
load  calculated  by the rational method of  FAR  23.425(d).  The 
downwash factor de/da is assumed to be 36a /AR^, per CAM  3.217(c) 
where a^ = slope of lift curve of wing (Cr per degree) and AR^  = 
aspect ratio of wing. Then the largest up and down tail loads are

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SELECTION OF CRITICAL LOADS

selected  as the critical up and down gust tail loads with  flaps 
extended.

Unsynunetrical Tail Load

The  computer program selects the maximum loading  from  the 
symmetrical flight conditions and applies 100 percent to one side 
of  the  plane of symmetry and lOO-lO(N-l) percent to  the  other 
side  per  FAR  23.427(b)(l)  and (2) where N  is  maneuver  load 
factor.

For  airplanes that are not conventional (such as  airplanes 
with  horizontal  tail surfaces having  appreciable  dihedral  or 
supported  by  the  vertical  tail  surfaces)  the  surfaces  and 
supporting  structures must be designed for combined vertical and 
horizontal  surface loads resulting from each  prescribed  flight 
condition taken separately according to FAR 23.427(c).

VERTICAL TAIL LOADS

Sudden Full Rudder Deflection

The  vertical tail side load for sudden displacement to  the 
maximum  deflection  is  calculated at V^ with  the  airplane  in 
unaccelerated flight at zero yaw as required by FAR 23.441(a)(1). 
The  load need not exceed pilot effort, but the computer  program 
does  not account for this and calculates the load for  full  de- 
flection.

The  load is calculated using the average loading of  B23.11 
and the figure Bl of Appendix B and the distribution in figure B7 
per FAR 23.441(b).

Yaw to Sideslip

The  vertical  side tail load for rudder deflected  to  full 
deflection  and airplane yawed to sideslip angle of 19.5  degrees 
is  calculated as required by FAR 23.441(a)(2) using the  average 
loading  of B23.11 and figure Bl of Appendix B and the  distribu- 
tion in figure B6 of Appendix B per FAR 23.441(3).

Yaw to 15 Degrees with Rudder in Neutral

The  vertical  side tail load for a yaw angle of 15  degrees 
with  the  rudder control maintained in the neutral  position  is 
calculated  as  required by FAR 23.441(a)(3)  using  the  average 
loading   of  B23.11  and  figure  Bl  of  Appendix  B  and   the 
distribution in figure B8 of Appendix B.

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FAR 23 LOADS

Lateral (Side) Gust Load

The lateral gust load in unaccelerated flight at V^ required 
by  FAR 23.443 is calculated using the rational analysis  of  FAR 
23.443(b).

FUSELAGE LOADS

Fuselage Vertical Shear

The  fuselage  is  critical for vertical shear  aft  of  and 
adjacent  to the rear spar attachment resulting from the  maximum 
net  up  load on the wing. It may also be critical  for  fuselage 
vertical  shear forward of the wing forward attachment. The  com- 
puter program searches the V-n data for the largest wing up  load 
accounting  for relieving wing inertia. For aft fuselage  mounted 
engines  this condition could also be critical for  aft  fuselage 
bending.

Aft Fuselage Down Bending

The  aft  fuselage is critical for down bending due  to  un- 
checked pull-up maneuver. See the discussion of FAA policy  under 
critical tail loads for unchecked pull-up maneuver.

The  aft fuselage is critical for the largest  down  bending 
from the combination of down tail load and down fuselage  inertia 
in  balanced flight conditions. The method used is to search  the 
balanced  V-n data for the largest product of down tail load  and 
fuselage down load reacted by the wing accounting for   relieving 
wing inertia.

Forward Fuselage Down Bending

The forward fuselage is usually critical for the same condi- 
tion  as  maximum aft fuselage down bending since it  reacts  the 
critical  aft  fuselage down bending moment with  limited  relief 
from wing torsion.

The forward fuselage is critical for down bending in 2 wheel 
landing  conditions,  but  is  usually not critical  compared  to 
flight  loads.   It  is  easily  written  off  by  comparing  the 
unbalanced moment in landing to the flight bending moment at  the 
forward wing attach point.

Fuselage Up Bending in Flight

The fuselage up bending moment in balanced flight conditions

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SELECTION OF CRITICAL LOADS

are  about  40  percent of down bending moment.  The  up bending

strength  is  usually more than 75 percent of  the  down bending

strength  of  the  fuselage. Therefore the  fuselage  up bending 
moment is usually not critical for flight conditions.

The fuselage up bending loads in accelerated pitching maneu- 
vers  are discussed in horizontal tail loads. The up load on  the 
horizontal  tail during the unchecked push down maneuver  is  not 
critical  for  fuselage up bending since it is in  opposition  to 
linear and pitching inertia.

Fuselage Up Bending in Landing

The  forward fuselage is critical for 3 wheel level  landing 
at either max landing at forward cg or most forward cg regardless 
of weight.

CHANGES IN REGULATIONS

Amendment  42 removes Appendix B from the  regulations.

For  ultralights, kitplanes and experimental homebuilts  the 
tail loads calculated with the Appendix B are very acceptable for 
safe  flight.  Thousands  of airplanes built  by  Piper,  Cessna, 
Beech,  Aero Commander the last three decades have been  designed 
and  substantiated  for strength to Appendix B.  They  are  still 
approved as airworthy by the FAA.

Rational  tail loads will be required  for airplanes to  be 
type  certificated.  In the near future  a  supplemental  program 
SELECT42.BAS  will be available which replaces the 5  tail  loads 
calculated with Appendix B with rational loads.

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