Technology Notebook Contents

Flight Loads

This section presents data development methods pertaining to the Airloads Survey Tests and Flight Loads Demonstration Tests, Landing and Ground Handling Tests, Wing Pressure Survey Tests, Aerial Delivery Subsystem Tests, Off Runway Operations and airframe Dynamic Response Tests. The methods presented herein generally have been proven on actual flight articles represent current state of the art techniques. The sketch in Sketch 1 presents an over all picture of the load components.

LOAD MEASUREMENTS

The structural load measurements for the airload survey and demonstration, the ground loads tests, the off-runway tests, dynamic response tests, and aerial delivery tests are generally obtained by the use of a strain gage system. Loads and air vehicle measurements are recorded on magnetic tape and normally employee real time telemetry monitoring of all maneuvers during the buildup and demonstrations.

Although a temperature compensation gage installation results in a constant slope of the gage output, temperature compensation cannot account for a structural zero load point shift due to changes in temperature. For example, a structure may be exposed to a constant load in a temperature varing environment. As the temperature varies, the structure will deflect differently and hence the strain gage output will vary, again, even though the load is constant. This is illustrated in Figure 1. which is Strain Gage Output (in mv) at Various Applied Loads.

NET INERTIAL LOAD

Thus, the problem of determining a true zero load point becomes a paramount purpose of the airload survey. To facilitate discussion, the NET structural load is broken down in Figure 2 as follows.

Unit Structural Inertia Loads

The unit structural inertia loads are developed by weighing the parts of the aircraft outboard of the station when the strain gage measurement is to be taken. Hence, the weight outboard is Shear Inertia. If Bending is required, then a series of elements times the respective moment is summed to produce a Bending Inertia value. Torsion is obtained In the same manner but by summing about the Elastic Axis near the center of the structure. The sketch in Figure 3. illustrates this procedure.

Unit Fuel Inertia Loads

The unit fuel Inertia is are obtained by a similar method as the structural inertia. Here, however, a complete calculation Is performed for each Increment in fuel loading for what would be a normal fuel 'burn-off. The sketch in Figure 4a. shows the concept.

As stated above, for each element of fuel, an inertial value is computed and presented in a chart with one line for each tank and a chart for each station measures and shown in Figure 4b .

AIRLOAD COMPONENTS

As stated above. for one element of fuel, an inertia value is computed and presented in a chart with one line for each tank and a chart for each station to be measured.

Basic Airload --- The basic airload is the airload distribution over the structure when the net air vehicle lift is zero, i. e. in the case of the wing, it is the airload distribution at the angle of attack of zero air vehicle lift.

Additional Airload --- The additional airload is the variation in Airload distribution over a structure at the lift on the structure varies, i. e. for the wing, this is simply increasing/decreasing load due to increasing or decreasing angle of attack.

Together with the above definitions, it is assumed that zero load occurs at zero q and zero g. Furthermore, ft is assumed that of such effects as dynamic pressure, Mach number, fuel inertia, airplane configuration, etc., dynamic pressure ;has the greatest effect on basic Airload. It is further assumed that once the effect of dynamics pressure has been determined, the effects of the other variables can be determined.

Thus, the problem of determining a true zero load point becomes a paramount purpose of the survey. To facilitate discussion, the NET structural load Is broken down as follows.

The section of the flight envelope where dynamic pressure effects are the primary variable is investigated first normally. For the initial tests, this will be approximately 10,000 ft pressure altitude and speeds up to Vh in the clean configuration. The data to develop the basic airload is derived from trim, turn, and pushdown maneuvers conducted at the above mentioned conditions. The test data is illustrated Figure 5 for a constant q point.

TRIMS, TURNS,and PUSH-DOWNS

Point A) is a stabilized trim at 1.0g. Points B) through E) are stabilized trims at approximately .25, .50, .75, 1.0, of .80 of (Nlimit - 1)g's with Point F) a pushdown to approximate 0 g's (Alternately, smooth "roller coasters" may be acceptable if the airspeed / altitude variations are small.)

The incremental loads from each maneuver are plotted versus normal force coefficient, Cn as illustrated in Figure 6

where

1. Cn = (Nz * W)/qS

2. q = (Ve * Ve)/295. - psf

The above process is repeated at each q point being investigated. It is to be noted that loads for all q points are calculated incrementally from the same reference point. The reference point is usually a convenient trim point such as 200 KCAS, and Nz = 1.0 g at the test altitude. Once all q points are flown, the load intercept points (incremental load at Cn = NzW/qS = 0) are plotted versus dynamic pressure as shown in the sketch of Figure 7 .

As shown in the Figure 7 sketch, the loads are extrapolated to zero q. Thus, the true zero load point is established, i. e., the load at Nz = 0, q = 0 to coincide with the true zero load point then results in plots which define the variation of basic airload with dynamic pressure. The final form of the curve is illustrated in Figure 8 which is Basic Airload as a function of Dynamic Pressure, q - psf.

ENVELOPE EXPANSION

Once this baseline basic airload variation is determined, it can readily be expanded to include the entire flight envelope and include effects of Mach number, fuel inertia, airplane configuration, etc. This is accomplished by flying these conditions and reducing the resultant data as previously described. For example, to determine Mach number effects, a trim point at a specified q and Mach number previously obtained during the baseline will be used as an initial test point and expanded by data obtained at additional Mach numbers.

The variation of the additional airload is developed in a manner similar to that used for development of the basic airload. Namely, the variation of additional airload versus q is is first determined and then the data are expanded to cover influences such as Mach number, configuration, etc. Before beginning the discussion of additional airload determination it is worthwhile to re-emphasize the following facts.

1) Additional airload is defined as the variation of airload with lift

2) Lift is equal to Nz * W

3) The slope of air vehicle load versus load factor is generally constant for a temperature compensated strain gage.

Thus a plot of incremental load versus load factor is developed from the plots of incremental load versus Cn used to establish the basic airload variation. The sketch in Figure 9 . illustrates this initial plot.

As evident in the above sketch, the slope (L/N)inertia must be taken out of the slope (L/N)net to arrive at the airload slope (L/N)a. The (L/N)a divided be the test gross weight results in the unit additional airload, i.e., the variation of airload with lift, (L/NW). The unit additional airload is plotted versus dynamic pressure as illustrated in Figure 10 which is (L/NW) as a function of Dynamic Pressure, q - psf

Once unit additional airload curves similar to the above sketch have been obtained, the data are then expanded to include the effects of other variables such as Mach number, configuration, etc.

The preceding analysis methods are employed to develop basic and additional airload curves for:

1) swept axis shear, bending moment, and torsional moment loads for the wing and horizontal stabilizer,

2) vertical bending on the forward and aft fuselage,

3) axial loads in vertical controls such as the pitch trim actuator,

4) pitching moments on conventional pylons.

Although the preceding analysis produces basic and unit additional airload data for the horizontal stabilizer and pitch trim actuator, a more detailed analysis may be required to establish the complete horizontal stabilizer airload distribution. For example, the detailed analysis of horizontal stabilizer/pitch trim loads should require data to be obtained at constant elevator position and data obtained at consistent stabilizer incidence angle. First, additional airload plots versus q are developed from data obtained with a fixed horizontal stabilizer incidence angle ( that angle required for the particular test speed, cg, etc. ). Load factor is obtained by variation in elevator position. In addition, the variation in incremental load due to elevator deflection is obtained. The sketches, Figure 11 are representative of the data, where ih = constant and de = variable.

In addition, basic and unit additional airload variations are developed for a constant elevator position with zero stick force. Variation in load factor are obtained by varying the horizontal stabilizer incidence angle. The data plots are illustrated in Figure 12 for Fe = 0 and de = constant.

Pitch trim actuator loads are handled in the same manner as described above for the horizontal stabilizer. Plots of basic and unit additional airload variations versus q for constant de and zero stick force, and plots of unit additional airload variation in incremental load due to elevator position versus q were developed. This document details how the above data is employed for actual load extrapolation.

SIDESLIP LOADS

Once the basic and additional airload curves have been established, the emphasis of the airload survey switches to determining incremental structural loads due to steady yaws, dynamic yaws, and aileron roll.

Trimmed zero yaw flight conditions are taken as a zero load point reference and incremental fuselage and vertical stabilizer loads are developed. Zero for rolling moment and side loads for the pylons are obtained by extrapolation of trimmed flight loads, (B=0) to a B=0, q=0, and Ny=0 condition. Incremental loads versus steady sideslip angle are plotted as shown in the Figure 13 sketch.

The sideslip angle is limited to that angle which results in maximum rudder hinge moment, rudder pedal q stops, or 80% structural loads at the test q. If the 80% limitation is reached prior to reaching hinge moment or pedal stops, the incremental loads are extrapolated to the steady state sideslip angle at maximum hinge moment or rudder pedal stops as illustrated above.

In addition to the steady state yaws, abrupt rudder kicks will be conducted to investigate the effects of any dynamic overswing during the maneuver. It should be noted that the directional control systems are generally designed for full time operative yaw damper and mechanical rudder pedal q stops. With regard to the yaw damper, it is quite possible, even probable at high q, that there will be no dynamic overswing during an abrupt rudder kick. The abrupt rudder kicks should be accomplished in the buildup phase to either maximum hinge moment, maximum rudder deflection, or 80% loads, whichever occurs first. a typical plot of the incremental loads including the effects of overswing is illustrated by the Figure 14 sketch for Ve = constant.

As with the basic and additional airload data, sufficient yaw maneuver data are obtained to define the effects of other variables such as gross weight, Mach number, configuration, etc.

AILERON (SPOILER) ROLLS

Incremental wing, aft fuselage, and horizontal stabilizer loads due roll maneuvers are determined from aileron rolls conducted through 45 deg of roll right and left maintaining as constant a load factor through the roll as possible. The loads are determined incrementally from the steady turn trim point prior to aileron deflection. Each series of maneuvers is conducted up to 100% total aileron deflection or 80% structural loads, whichever occurs first. The incremental loads at maximum roll rate and maximum roll acceleration are plotted versus percentage total aileron deflection as shown in the Figure 15 below.

In the event 80% load limitations are reached, the incremental loads are extrapolated to 100% total aileron at the test Q. The loads at 100% total aileron are plotted versus q as indicated in Figure 16 .

As shown by past experience, the preceding analysis does not necessarily define the peak incremental load occurring during the run. Therefore a separated analysis is conducted to determine the maximum incremental loads anytime during the run. For example, consider the wing loads shown in the Figure 17 sketch.

From a series of runs at constant q similar to the above sketch, peak load plots are constructed as in Figure 18 from which the following are obtained.

The loads at 100% total aileron are them plotted versus q to show the incremental load variation over the q range. The load plots versus q are then evaluated to define the critical load combination with regard to design load envelopes.

TEST REQUIREMENTS

Airload Survey

Airloads survey data analysis is based upon the fact that a strain gage gives a measure of NET load acting on a structure under any flight or ground condition. The strain gages used in test are normally calibrated and temperature compensated in the installation. Water proofing of the gages is essential. Since the gage installations are temperature compensated, the strain gage output versus load is linear and has a constant slope regardless of the temperature environment in which the gage is operating.

Prior to conduction the 80% and 100% flight demonstration maneuvers, the airload survey loads data are extrapolated to the flight conditions planned for conducting these maneuvers. The extrapolated data serve to,

(1) Verify the analytical determinated criticality of the maneuver as planned, and

(2) define possible new critical maneuver conditions not presently defined analytically.

Once the maneuver conditions are defined in terms of q, Mach, configuration, gross weight, fuel distribution, load factor, etc., the required extrapolation can be readily accomplished.

Trim loads determined from basic and additional airload calculations are added to the appropriate incremental maneuvers determined from the steady state yaw and roll data. For example, consider an aileron roll maneuver. The trim loads prior to initiation of the roll are calculated from the following Equation (1).

Lt = [ L/NW x Wm + (L/N)s + (L/N)f ] x Nm + Lb................................( Equation 1 )

where

Lt = trim load

Wm = maneuver gross weight

(L/NW) = unit additional airload

(L/N)s = unit structural inertia

(L/N)f = unit fuel inertia

Nm = trim load factor

Lb = basic airload

It is to be noted that (L/N)f in the above equation is applicable only to wing loads. The term in equation (1) are determined from the plotted airload survey data as illustrated in Figure 19a and in Figure 19b for a sample wing load.

The incremental load due to the aileron input is determined from a similar plot as shown Figure 20 .

The extrapolated net maneuver load is simply the summation of the trim load and the incremental load in Equation 2.

Lnet = Lt + dLm ................................ ( Equation 2 )

In a manner similar to equations (1) and (2), horizontal stabilizer loads and pitch trim actuator loads can be determined for any load factor condition as shown in equation (3) below.

Net Load @ Nzm = (Lb)@Fe=0 + [(L/NW)@Fe=0 x Nz2 x Wtest] + [(L/N)i X Nzm ] + [(dL/dELEV) x ELEV] + [(L/NW)@dELEV=var x Nz x Wtest ] ................................ ( Equation 3 )

where

Nzm = Target load factor for maneuver

Lb@Fe=0 = Basic airload for zero stick force

(L/NW)@Fe=0 = Unit additional airload for zero stick force

Nz2 = 1.0

Wtest = Test gross weight

(L/N)i = Inertial load

dL/dELEV = Load variation due to elevator

ELEV = Elevator position for 1.0g trim @ test conditions

(L/NW)dELEV=var = Unit additional airload for variable elevator

Nz = Nzm - 1.0g

During actual conduct of an 80% and 100% maneuver, a trim record is obtained prior to conducting the maneuver. The trim conditions are of course known and hence the trim loads are readily calculated from basic and unit additional airload data previously. The incremental maneuver load is then added directly to the trim load resulting in the maneuver.

Flight Demonstration

(to be added)

80% Demonstration

(to be added)

100% Demonstration

(to be added)

Static Test

(to be added)

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