STRUCTURAL DESIGN PHILOSOPHY - ICAS 353 Heppe et al.pdfExisting discret e gust criteria are known to provide satisfactory structure for aircraft similar in configuration, flight profile,

STRUCTURAL DESIGN PHILOSOPHY

FOR THE SUPERSONIC TRANSPORT

R. RICHARD HEPPE

Chief Advanced Systems Research Engineer

M. A. MEIXON

Department Manager, Structural Methods

W. A. STAUFFER

Department Manager, Basic Loads

Lockheed-California Company Burbank, California

INTRODUCTION

The Lockheed Aircraft Corporation is currently participating in theNational Supersonic Transport Program under study and development int he United States. In response to t he original Request for Proposal fromt he Federal Aviation Agency, Lockheed proposed in January 199-1 a fixed,double-delta-wing airplane designed around the fundamental goal of maxi-mum productivity and optimum airline economics for operation in thepost-I970 time period. The general arrangement of t his airplane is shownin Fig. 1. Thorough analysis of alternative means of at taining t he objectivesled to the combination of a large fuselage capable of housing up to 221passengers, a high cruising speed of Mach 3.0, and the large fixed wing. Itis t be purpose of this paper to deal in dept h with t he basic struct ural designphilosophy being applied to t his airpIane.

The recommendation of a high cruising Mach number was based uponfundamental considerations of optimum range, payload, and directoperating cost t ogether with a high Hight altitude tor minimum sonic boom.The Afach 3.0 speed is recommended wit h full understanding of theprogressively more hostile environment in t he high-Afach-number flightn'gime, a factor which has }Wen accounted for in t hese analyses, both with

353

FOURTH CONGRESS - AERONAUTICAL SCIENCES

7--. '":1Prrrrtr35

Figure 1. Lockheed Double Data supersonic transport.

respect to development time and cost. Materials of construction are beingselect ed in detail to meet the requirements of this environment whilemaim outing or improving the levels of safety and reliability which havebeen developed in modern transport operation. It is not the purpose of thispaper t o discuss detail design nor the detail selection of materials in indi-vidual areas, but it should be stated at the outset that the principalmaterials of construction will be of the titanium family in a number ofalloys as suit ed to individual applications.

To attain the economic, reliability, and safety objectives of the super-sonic transport, a well defined and proven structural design philosophy isbeing applied. This philosophy has evolved front more than 30 years ofdesign, production, and airline operation of transport aircraft throughoutthe world. Although the environment is new and new mat erials will beemployed, t he fundamental approach is the sante one which has contributedto successful airline operating aircraft over this time period.

The five principal elements of the supersonic transport structural designphilosophy are (1) the use of rational design load criteria, (2) the attainmentof unlimited service life, (3) the attainment of an airframe having fail-safecharacteristics with respect to both strength and stiffness, (4) recognitionfrom the outset of the role of inspection and maintenance in structuraldesign, and finally (5) an extensive ground and flight structural provingeffort. The remainder of this paper describes the approach being taken ineach of these five areas.

STRUCTURAL DESIGN PHILOSOPHY 355

STRUCTURAL CRITERIA

The basic structural design philosophy requires that criteria be respon-sive to the operating environment and novel characteristics of the super-sonic transport and, at the same time, provide a level of safety at leastequivalent to that of current transports. The structural airworthinessrequirements of the U. S. Civil Regulations (CAR-4b) are known to haveprovided satisfactory levels of strength for current civil transports. How-ever, in those flight regimes where the operating environment or theresponse characteristics of the supersonic transport differ markedly fromearlier transports, it is important that the adequacy of past criteria bereviewed. Some of the areas requiring such consideration are discussedbelow.

FLIGHT PROFILE

The first element of the rational approach to structural criteria for designof the supersonic transport involves the actual flight profile. Significantvariations in flight profile are anticipated for different operating ranges,routes, and because of air traffic control requirements. Nevertheless, theflight profile shown in Fig. 2 for a range of 3,500 nautical miles is repre-sentative and can be used to demonstrate critical aspects of the plannedoperation.

The supersonic transport is characterized by large weight changes dueto fuel burnoff during flight. For example, the gross weight may reduce asmuch as 70,000 lb from takeoff to the start of high-Mach-number cruise.During the cruising portion of the flight, over 100,000 lb of fuel may beburned off. These large variations in gross weight combined with the largechanges in Iach number, altitude, and environmental temperature suggest

ALTITUDE

0 1 2 3 4RANGE — 1000 NAUTICAL MILES

Figure 2. Typical design flight profile.

ALTITUDE

-1000

FT. 80

40

o

356 FOURTH CONGRESS — AERONAUTICAL SCIENCES

that rational combinations of weight, speed, and altitude should be con-sidered rather than arbitrary ones. Accordingly, structural design weightsare selected for specific altitude and flight conditions with due considerationbeing given to possible variations front these values for a variety ofoperating conditions.

DESIGN SPEEDS

The selection of design cruising, Vc, and dive, VD, speeds is made withconsideration of such factors as structural weight, flutter, sonic boomoverpressure, airplane performance, and inadvertent overspeed. For thesupersonic transport, the selection of these design speed values is a part icu-larly sensitive matter, and the values shown in Fig. 3 are but representativeof the factors involved.

At subsonic speeds, Vc must be high enough to permit economic climband cruise performance, but low enough to avoid excessive weight penalties

ALTITUDE

-1000

FEET

oo

ao

70

60

50

40 2.5 PSFSONIC BOOtA

30

20

10

„

-.9

vc/mc

360 KEAS.-

460 KEAS

585 KEAS

V0/.0

450 KEAS

0 100 200 300 400 500 600 700

VE - EQUIVALENT AIRSPEED - KNOTS

Figure 3. Design speed—altitude relationship.


due to gust loads. At intermediate altitudes, some increase in Vc leads todecreased climb fuel, but must be balanced against the weights involved indemonstrating aeroelastic and flut ter characteristics with an adequate divespeed margin. In addition, adequate margin must be provided for climband acceleration along different sonic boom paths so as not to compromisethe eventual operational flexibility of the airplane. Finally, at high altitude,the Vc becomes related to the design cruising Mach number and mustadequately cover the needs of the airplane to realize optimum long rangecruise capability.

Of particular importance and interest are the speed margins between thedesign cruising speed and/or Mach number values and the dive limits.Margins which are felt to be adequate for a variety of rational conditionsare shown in Fig. 3. Typical inadvertent overspeed conditions which havebeen investigated inelude a 7.5-degree nose-down upset from cruise con-ditions held for 20 sec, autopilot "hard-overs," trim system malfunctions,passenger movement, level off from climb with delay in power reduction,and level off at low altitudes while crossing airways with delay in powerreduction. For all of the conditions considered, it has been found that adive speed margin of 0.2 Mach number above the cruising speed is adequatefor cruise operation. At lower altitudes for various conditions during bothascent and descent, the minimum overspeed margin requirements are ofthe order of 100 knots equivalent airspeed.

TEMPERATURE CRITERIA

Steady-state and transient temperat ure conditions and the aerodynamiccharacteristics of the airplane appropriate to each particular flight, condi-tion are considered for all loading conditions. Both normal and emergencyconditions are considered in the flight profile analysis which determines 1:heappropriate structural temperatures and loads. To protect the airplanefront hotter-than-standard day conditions, a total temperature placard isestablished. For example, this placard, TAm, might be defined as 650°F,corresponding to Mach 3.0 under U. S. Standard atmosphere conditions at80,000 ft. This placard is shown in Fig. 4.

The structure is designed for fat igue, fail-safe, and ultimate conditionsfor the equilibrium and transient temperatures along the Ve//11c versusaltitude profile. The climb segment is conducted at maximum climb powerfor all gross weights from minimuni to maximum in order to produce ihemaximum heating rates and greatest thermally induced st resses. Loads dueo conditions of limit severity (pitch, roll, and yaw maneuvers, gust s, engine

failure, etc.) are combined with thermally induced st resses and tempera-tures at any point on the flight profile. For maximum rat es of cooling,


900

ALTITUDE 80,000 FEET

800

n 700

.1 600

ootl•

500

400

- 100 - 80 -60 -40 -20

AMBIENT TEMPERATURE - °F

Figure 4. Design temperature limits.

emergency descent along the Vc/Mc boundary is assumed from any pointalong the design flight profile.

In addition, the design provides for the transient temperatures and loadsresulting from inadvertent overspeeds to VD/MD. The initial condition isassumed to be flight along the Vc/Mc boundary with initial attitudes,power settings, weights, and temperatures appropriate to the particularpoint in the flight profile where the overspeed is assumed to occur. Timesto reach dive speed and to return to the Vc/Mc placard boundary are afunction of the type of upset condition. Shortest times to accelerate toM = 3.2 have been found to occur due to leveling-off from climb with lag inpower adjustment. The time at VD/MD is assumed to be 10 see and loadsof limit severity are assumed to occur during the overspeed.

The combination of load stress and thermal stress due to temperaturedifferentials is fundamentally simple. Thermal expansions in a staticallydeterminat e structure will create deformations without correspondingresisting loads. Only a redundant or restrained structure will developresisting loads. The total thernial expansion of elements is relaxed anddivided among surrounding elements according to the flexibility of thetotal structure. Thermal stresses are, 1 herefore, the result of, partiallyrelaxed strains in the redundant structure. There is no resultant external

MIL-STD-2

10A-CO

LD

U.S.

STD

ATM,

1962

MIL

STD-210A

HOT


load; the internal force systems at any cut section are entirely self-equili-brating. On the other hand, external load conditions provide the initiatingstresses and result in st rains in t he elements of the structure. The distinctionis important for the prediction of the ultimate strength of the structurewhere plasticity or yielding of the material must be account ed for rationally.

The approach currently in use for the supersonic transport is that thecombination of effects from external loads and temperature differentials bebased on the suns of strains rather than stresses. The design stress isobtained from the stress-strain curve for the appropriate teMperature,corresponding to the sum of these strains. The approach is graphicallyillustrated in Fig. 5. The specific criteria is as follows:

For ultimate tension or compression conditions, a factor of safety of1.25 is applied to calculat ed thermal strains when load and thermalstrains are additive. No factor is applied to thermal strains whenthey are relieving the load strain.

For fatigue-limited tension conditions and for fail-safe conditions, nofactor is applied to thermal strains.

STRESS-STRAIN CURVE FORDESIGN TEMPERATURE

ALLOWABLESTRESS

I .2 5E , „

to-Ef„

STRESS

DUE TO ULTIMATE LOAD= STRAIN

r,„ = THERMAL STRAIN

Figure 5. Thermal stress accountability—ultimate conditions.

360 FOURTH CONGRESS - AERONAUTICAL SCIENCES

GUST LOADS

Existing discret e gust criteria are known to provide satisfactory structurefor aircraft similar in configuration, flight profile, and design envelopeutilization to current aircraft. For an advanced vehicle such as the super-sonic transport , however, it is important that the adequacy of these criteriabe reviewed, utilizing the more realistic continuous random description ofatmospheric turbulence and analysis techniques now available, to assurethat a level of strength consistent, with that of past aircraft is maintained.Experience has shown it to be particularly important to utilize a continuousrandom turbulence description for gust-loads work if the damping in eitherthe airplane short-period modes, or in any of the elastic modes, is signifi-cantly different from that of airplanes whose operational experience wasused to establish current gust criteria.

To accomplish this purpose, design load levels for the wing, fuselage, andvertical tail are defined statistically by means of a power-spectral flightprofile analysis, using the general procedure and atmospheric model ofNACA TN 4332. The basic philosophy in this work is to provide a level ofstrength such t hat, per flight hour, the probability of exceeding limitstrength is no greater for the supersonic transport than for present com-mercial transports having proven adequacy for gust-induced loads. TheModel 749 Constellation is representative of such existing aircraft. Thesupersonic t ransport power-spectral flight profile analysis provides curvesof frequency of exceedance versus load level for any desired load quantity.Entering these supersonic transport curves at the limit load frequency ofexceedance established for the Model 749 then establishes the limit designstrength level required for supersonic transport loads.

Wing root-bending results for the supersonic transport are shown inFig. G. Statistical curves are shown separately for various portions of theflight profile. These separate curves, when added vertically, give the totalfrequency of exceedance. The most st riking result shown by this figure isthat t he supersonic t ransport wing will be exposed to far higher gust loadsduring subsonic climb than during the other portions of the flight.. As notedpreviously, however, the subsonic climb speed has been selected low enoughto prevent 1 he gust loads front exceeding the 2.5-g wing maneuver loads.To place in perspective the gust, criteria resulting from the flight-profileanalysis, it is of int erest to not e the discret e gust velocity (Ud,) that wouldbe necessary, on a static load basis, to provide the same wing-strengthlevel is 69 fps, as compared to the current. civil criteria value of 50 fps at Vc.

MANEUVER LOADS

The supersonic transport is designed to a positive maneuvering loadfactor of 2.5 at all speeds front the design maneuver speed (VA) to the


N(Y)

-EXCEEDANCE

CYCLES/1000

HR.

100000

10000

1000

100

10

SUBSONIC CLIMB

SUBSONIC ACCELERATION

SUPERSONIC CLIMB,CRUISE & DESCENT

SUBSONIC DESCENT

TOTAL

.010 20 40 60 80 100 120

WING ROOT BENDING MOMENT - MILLION IN. LBS.

Figure 6. l'ower spectral gust loads.

design dive speed (VD). The negative maneuvering load factor is —1.0 atall speeds up to Vc and varies linearly from — 1.0 at Vc to zero at VD. Thisis in accordance with current U. S. civil requirements. Typical speed-loadfactor diagrams at six altitudes and representative weights are shown inFig. 7. These diagrams are superimposed on the design speed-altitudediagram. At each altitude the diagrams show both the minimum and themaximum design weights at the particular altitude. The low-speed portionof these V-n diagrams is defined by a stick-shaker warning device and doesnot represent a physical stall speed. The warning device is incorporated tolimit angle of attack to appropriate values in the different operating speedregimes. Experience with low-aspect-ratio delta wings establishes the factthat no buffet-boundary limits are anticipated within the V-n diagramsshown. This capability provides added flexibility in selecting cruise alti-tudes to alleviate sonic boom, and freedom for the operator to selecttransonic acceleration altitudes.

The control system, inertia characteristics, aerodynamic characteristics,flight profile, and speed range of the supersonic transport are su fficientlydifferent from those of current transports to raise questions as to theadequacy of existing civil criteria for abrupt pitch, yaw, and roll maneuvers.Therefore, the time history response to rationally defined control inputsand/or engine failures are determined at appropriate altitudes to evaluatethe response of the airplane in the subsonic, transonic, and supersonic speedregimes for all points within the design V-n and weight-center-of-gravityenvelopes. Results of these analyses have shown that the supersonic

STRENGTF1 LEVEL FOREQUIVALENT SAFETYTO MODEL 749 7

3( i2 FOURTH CONGRESS — AERONAUTICAL SCIENCES

transport is not dependent on stability augmentation to keep the designloads for transient roll, pitch, yaw, and engine failure conditions to accept-able levels.

LANDING CRITERIA

In order to obtain a preliminary insight into the adequacy of currentcivil aircraft standards for design sinking speed of the supersonic transport,use has been made of two parameters suggested in Part II of NASA TND-1392 as possible explanations for the difference in operational sinking-speed statistics of subsonic jets and piston aircraft. These are an elevatoreffectiveness parameter, and a parameter roughly measuring the pilot'sheight above the ground at main gear contact. Values of the sinking speedexceeded once in one hundred landings are plotted versus these parameters

90 — — — MIN WT

212000 LB I

ALL ALT _

BO

70

3

2

321 n

30

20

244100C!

_0 100 200 300 400 500 600

EQUIVALENT AIRSPEED - KNOTS

Figure 7. Typical design (V-n) and weight relationship.


for various existing aircraft on Figs. 8 and 9 with values for the supersonictransport superimposed. These results indicate that supersonic transportsinking speeds are likely to be essentially the same as for current subsonicjets. It should be noted that, even if supersonic transport sinking-speedstatistics should prove to be greater than those of current jets, it is notapparent that the design level of 10 fps should be raised. This is becausecurrent jet statistics show that a sinking speed of 7.5 fps is exceeded onlyonce in 10,000 landings. Supersonic transport sinking-speed statistics couldbe as much as 33 percent above current jet sinking speeds, yet the sinkingspeeds reached once in 10,000 landings would still be only 10 fps. Thisexceedance rate would not appear to be unreasonable for limit design.

A nose landing gear criteria which reflects the airplane's pitch-plungemotion during main and nose gear impact has been employed. The criteriaassumes main gear impact at the limit sink rate of 10 fps in the maximumnose-up attitude, most forward center of gravity, and further assumes nopilot action during the contact of the main gear and settling of the nosegear on to the runway. Figure 10 shows the results of such an analysis. Inthe example shown, it can be seen that the sinking speed of the nose gearis 10 fps at main gear contact, reducing almost to zero as the main geararrests the vertical descent of the airplane, and then gradually increasingas the airplane rotates until the nose gear contacts at a sinking speed of7 fps. It is of interest that the resulting nose gear vertical load is less thanhalf that resulting from dynamic taxi.

TAXI CRITERIA

The taxi loads for the supersonic transport are based upon provencriteria and experience with subsonic aircraft. The supersonic transportwill, of necessity, have to operate in the same general environment as

PISTON

4 6 8

V3CM8e CLa

lip.2Ky2

VERTICAL

VELOCITY

FT/SEC4

2

TURBOJET

SST * •ELECTRA

2

Figure 8. Elevator effectiveness—landing rate of sink relationship.

FOURTH CONGRESS — AERONAUTICAL SCIENCES

subsonic jets during takeoff, landing, and taxiing on rough runways ortaxiways. There is no significant difference in landing or takeoff speeds;and, because taxiways are generally rougher than runways, the taxi speedon taxiways will continue to be limited by the pilot's judgment based on

6

TURBOJET

VERTICAL

VELOCITY

- 4

2

O

SST

ELECTRA

PISTON

04008001200

W L

CL.Sq

Figure 9. Effect on landing rate of sink of pilot's eye height.

10AIRPLANE PITCH ATTITUDE - DEG.

0

200,000

100,000

MAIN GEAR VERTICAL LOAD - LBS.

o

o5 MAIN GEAR STRUT POSITION - IN.

10

10

5 NOSE GEAR SINKING SPEED FT/SEC

0

100,000

NOSE GEAR VERTICAL LOAD - LBS.

O

0

5

10

15NOSE GEAR STRUT POSITION - IN.

o 2345

TIME, SEC.

Figure 10. Landing time history.


roughness of the ride. Assurance of the validity of applying criteria devel-oped for subsonic aircraft to the supersonic transport has been obtained byanalog analyses in which the supersonic transport is taxied over knownrunway profiles and compared to existing aircraft. Pilot compartmentaccelerations determined in this manner for the supersonic transport arecomparable to those for subsonic airplanes and are not great enough tocause difficulty to the pilot in the takeoff run or to limit taxiway speedsbelow those of existing aircraft.

FLUTTER, DIVERGENCE, AND CONTROL CRITERIA

The design approach to flutter safety is based on the requirements inCivil Aeronautics Manual 4b with the objective of assuring that the airplaneis free from flutter of wing, vertical fin, and control surfaces, and free frompanel flutter at all speeds up to 1.2 VD. Fail-safe criteria include flutterconsiderations in that no single failure will cause flutter at any speed upto VD. All parts of the airplane will be demonstrated in flight to be freefrom flutter and excessive vibration under all speed and power conditionsappropriate to the operation of the airplane up to at least VD.

The design speed envelope previously described contains four regionswhich may potentially design the airplane or some of its elements forflutter and aeroelastic considerations. At low altitude, with the aerodynamiccenter located most forward, certain flutter modes may be critical. Undertransonic conditions, maximum values of lift curve slope occur indicatinganother potentially critical area. The other two potentially critical con-ditions occur along the dive-speed boundary under conditions of maximumdynamic pressure but with various conditions of lift curve slope andstructural temperature. In addition to these potentially critical areas offlight, certain configuration features, such as the location of the wing onthe fuselage and the engines on the wing, present t he possibility of differentmodes than are common in transport aircraft today.

An extensive analytical program is being augmented by elastic modelt esting to give precise definition to flutter limits for both normal conditionsand for conditions after single element failure. To date, analysis and testingindicate that flutter prevention on the double-delta configuration does notintroduce unusual structural stiffness requirements or complications, andt he anticipated development program including event ual flight flutterdenionst rations will positively est ablish t he airframe int egrity.

UNLIMITED SERVICE LIFE

After the selection of rational load crit eria based upon t he int ended usageand environment , t he next eletnent of 1 he structural design philosophy for


the supersonic transport involves operational life. Here t he design objectiveis to provide an airframe which will experience essentially unlimited life innormal airline duty. In specific terms, the phrase "unlimited life" meansthat with normal inspection, maintenance, and repair the operational lifeof major structural elements will be determined by inevitable economicobsolescence factors of the airframe and not by any requirement to retiresuch elentents on the basis of accumulated hours. Not only is this structuraldesign approach similar to that which has been used for the design of mostcurrent airline aircraft, but it also is required to substantiate the antici-pated depreciation write-off periods visualized as fundamental to theeconomic operation of such an airplane. For example, operation of theairframe for 3,000 hr each year for a period of 15 years will accumulate45,000 hr of service time. Such service lives are already relatively common-place with numerous Lockheed Constellation airframes now nearing 60,000hr of service on the airlines without replacement of primary structuralelements. Achievement of a similar service life for the supersonic transportis essential and is adopted as an element of the structural design philosophy.

CREEP CONSIDERATIONS

Perhaps the first requirement for achievement of unlimited service lifeis to insure freedom from significant creep of materials under the load andtemperature environments involved throughout the airframe st ructure. Forthe supersonic transport, similar criteria have been adopted as for previoustransport aircraft. For example, there will be no requirement to replacemajor structural elements at specified time intervals, nor to retire theairframe at any arbitrarily defined lifetime.

Control of creep depends upon the proper choice of material for theintended operating environment, together with control of the permissiblestress levels under operating conditions. Test results are already availableto guide the selection of materials. Typical long-time, high-temperaturetest data on a number of materials are shown in Fig. 11. The creep stress-to-density ratio information shown indicates that the titanium alloy8AL-1 Mo-1V has a clear superiority over other titanium alloys andstainless steel. Inconel alloys have even greater creep resistance at highertemperatures, but the material density precludes its general use on thesupersonic transport. Other alloys of titanium also produce excellent creepcharacteristics in the temperature range from 500 to 800°F, and comparisonof the capability of 8-1-1 titanium with aluminum at much lower tempera-tures shows that, use of proper mat erials can provide equal or less creep inthe structure of a Mach 3 supersonic transport than can be expected at,lower Mach numbers with aluminum.


A specific example of the application of titanium 8-1-1 to a critical hotelement is shown in Fig. 12, which presents information for the engine inletduct. In this area, short exposures to temperatures above 600°F are antici-pated, but a very large total of a total service exposure of 50,000 hr, halfor more, might be at temperatures near 600°F and at a stress level of theorder of 40,000 psi. The figure shows the cumulative distribution of hoursat stress and temperature levels for this critical component, and comparesit with the stress levels producing 0.1 percent creep deformation afterlengthy exposure. The figure demonstrates that substantial design marginscan be provided, and significant creep can be eliminated in the supersonictransport airframe even in the hottest elements.

FATIGUE CONSIDERATIONS

With freedom from creep insured by the proper selection of materials andstress levels for each environmental area, the important. considerations forunlimited service life become those of fatigue. To provide a service lifeunlimited by fatigue considerations, the supersonic transport will bedesigned by the same procedures which have been developed over the yearson the Constellation series, the Electra, the P-3A, and the supersonic F-104.The fundamental steps are as follows:

A fatigue quality level is established. This quality standard is slightlymore severe than a simple open hole in the uniform structure wellremoved from jcints or discontinuities, and all elements and jointsare required to meet or better this standard.Reduced design stress levels are utilized. When the quality standarddescribed above is applied, ultimate tensile strength levels cannot beexploited, and reduced design allowable tension stresses are estab-lished by analysis and test..Parts designed to meet the fatigue quality standard must be provedby test. All vital joints and structural discontinuities are required todemonst rate acceptable fat igue life in laboratory tests conducted oncomplet ely detailed specimens and using t he full spectrum of expectedservice loads.Estimated loads are confirmed by flight-load surveys. Direct flightt est measurement of loads and t emperatures are made to correlatewith estimated loads and to provide a basis for complete airframefat igue t est ing.Full-scale airframe fatigue tests will he performed. These tests willbe based upon load spectra verified by flight test, and will be per-formed on a flight-by-flight basis including t he fatigue-sensitiveground-air-ground cycle.


STRESS TO NODUCE 0.1% CREEP IN 10 HRS.

STRESS 70 PRODUCE 0.1* CREEP IN 1030 MRS.

STRESS TO PRODUCE 0.1% CREEP IN 10,030 IIRS.

STRESS TO PRODUCE 0.1% CREEP IN 50,030 HRS.

EXTRAPOLATED0.7

0.6

2;34;0 7,6 ALUMINUM(

e • 0.100

0.5

R-I.1 TITANIUM

(F),, • 140 ROI AT R.T.)P 0.158

OA

4-3-1 TITANIUM

(190 RSI)P. 0.163

2618-76

E 0.3

0.2

0 .

2618-76 ALUMINUM

(58 000

Pll 15-7 .

(225 KS0E. 0.272

ss \ IPCONEL718 (215 RSOa • 0.297

s

100 200 X10 400 500 600 703 800 920 1000 1100 200

TEMPERAfURE

Figure 11. Creep strength to density ratio for sheet alloys.

This approach to design for long fatigue life results in a comprehensivetreatment of the airframe element by element, with redesign and retestrequired where unanticipated failures are encountered.

To achieve adequate fatigue life under the approach described, a con-ventional approach to the prediction of fatigue effects is utilized, butimportant differences are introduced in the interpretation of these data.The conventional approach -to the prediction of service life assumes thatfatigue effects under different loading conditions are linearly cumulativeand can be based upon a summation of conditions, each one of which isbased upon 8-N fatigue characteristics developed front constant-load-amplitude test data. These results produce estimates of service life forconditions resulting from various operational usage in terms of the ultimate


design tensile stress and a stress concentration factor or fatigue qualityindex. To correct for the inadequacies of cumulative damage theory and thequestionable effects of the ground-to-air cycle, flight-by-flight fatigue testdata are utilized to select the proper fatigue quality index K for the

airplane.To accomplish such estimates of fatigue life, various flight profiles

representing the anticipated actual applications of the aircraft to transportservice are analyzed. Such important factors as short- and long-range flightsand the use of the airplane on training flights where it ordinarily experiencessomewhat more severe loadings than in normal usage are thus included.Finally, summations of the total life of the airplane involving differentmixes of the anticipated flight profiles can be made to determine actuallife estimates.

A typical example of the results of the conventional portion of thefatigue life analysis is shown in Fig. 13. The anticipated service life for atypical structural element such as an 8-1-1 titanium lower wing-spar capis shown as a function of the ultimate tensile stress and fatigue qualityindex. Similar data are developed for the entire spectrum of the flightprofiles anticipated and for other representative primary structural ele-ments of the airplane. The question remaining to complete the specific lifeestimate is that involving the selection of a specific combination of stressand fMigue quality index.

To determine the proper fatigue quality index, flight-by-flight tests areperformed on specimens wit h a simple hole and utilizing a loading sequencerepresentative of actual use such as shown in Fig. 14. The results of these

90

aoSTRESS TO PRODUCE 1% CREEP

690°F DEFORMATION TITANIUM 15H70 DUPLEX ANNEALING

60

50

STRESS -KSI40

700°F610°F

610°F

CCUMULATED 11ME AT ORBOVE THE GIVEN STRESS FOR,OCO HOURS SERVICE

zo

10

01 5 10 20

TIME - 1000 HRS.

Figure 12. Engine inlet duct—accumulated stress temperature history in 50,000hours of service.


120

7- 3500 NAUTICALMILE FLIGHTPROFILE

110

K = 4.55 K = 4 K = 2.330— 100

:5 K = 5.5K = 5

O

3 9°O

813

o=5'

70

Figure 13. Calculated service life-3500 N. Mi. flight profile.

tests are analyzed by the simple cumulative damage hypothesis in combi-nation with S-N data obtained from constant-amplitude tests. Resultsshow that accumulation of fatigue damage in act ual sequential loadings,including ground-air-ground effects, results in allowable stress levels lessthan 1,hose predicted from the constant amplitude data. An example of thisis shown in Fig. 15, where a fatigue quality index of 4 for coupons withsimple holes has been found to be representative of charact eristics resultingfront the actual flight-by-flight load sequences. This is significantly moresevere t han 1 he data resulting front constant load-amplit ude tests onsimple hole fatigue specimens resulting in a value of K of 2.76.

From testing of this type on the materials of int,erest for the supersonictransport , a quality index based upon a simple hole in material removedfront a joint area would have a minimum value of K = 4.0. As indicatedin 1:he design approach 1,o fatigue life outlined above, the quality indexadopted for design is intentionally selected to be somewhat more severethan the simple hole, and an index of 5 is expected to be both realistic andattainable in act,ual airframe structure. Therefore, this quality index isused as a basis for the st met ural life est imat es.

Results of typical service life calculations utilizing t his approach areshown in Fig. 16. Various flight, profiles, including pilot training and check,are considered, and estimates are made for various anticipated combina-tions of the flights. An ultimate design gross area tensile stress of 90,000psi for this particular material was selected in conjunction with the fatigue


quality index K = 5 as a basis for the life estimates. The results shownpredict essentially unlimited life for all of the composite flight spectraexamined. These results provide justification for the selection of an ultimatedesign tensile stress of 90,000 psi for the wing structure in 8-1-1 titanium,and also indicate a basis for unlimited fatigue life of the actual airframe atthis stress level.

To provide further substantiation of the fatigue characteristics of theairframe under actual operation, extensive research programs under jointFAA-Lockheed sponsorship on the fatigue characteristics of duplexannealed 8-1-1 titanium, precipitation hardening 14-8 steel, and nickel alloyInco 718 are in progress. Over 6,000 specimens are being evaluated undervarious conditions of temperature and stress exposure, and the effects ofvarious types of contaminants on these materials are being determined. Inaddition, specimens are being subjected to real-time flight-by-flight testingin which the time at temperature during each flight is approximately onehour.

1.45 fig ref.

THERMAL STRESSINCREMENT

-41g ref .. ---1— 0.8 fig rei.

0.20 fig ref.

CLIMB CRUISE

ONE FLIGHT

DESCENT—. TAXI

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Figure 14. Flight-by-Hight loading sequence.100

Tr 8-1-1 DUPLEX ANNEALED

ROOM TEMPERATURE

fuLAN 25,C00 PSI

SIMPLE HOLE - CONSTANT AMPUTLOE TESTS

SIMPLE HOLE - FLIGHT BY FLIGHT TESTS

K, • 2.76

_

4.00

SD

60

-KS!

40 - -

'102 103 ,04 1115 ,06107

CYCLES TO FAILURE

Figure 15. Typical constant amplitude fatigue data.

372 FOURTHCONGRESS — AERONAUTICAL SCIENCES

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Figure 16. Fatigue life estimate for composite service use of supersonic transport.

FAIL-SAFE DESIGN

The objective of the fail-safe crit eria is to insure the st ructural integrityof the airframe in the event of the failure of a single primary structuralelement: or the occurrence of partial damage in extensive structure such asskin panels. Such failures or damage may arise from unreported accidentalimpact, minor collisions, turbine-blade penetration, small-arms fire, orother sources as well as from cracks initiating from fatigue. Under suchconditions, integrity of the st ructure must be maintained for both strengthand stiffness requirements. Fail-safe strength and stiffness are provided atlimit-design load level in all normal design flight, takeoff, and landingconditions. These levels exceed current U. S. Civil Air Regulations require-ments by approximately 25 percent. The approach to achieving the fail-safeobjective is through the design process in which careful attention is givento the application of redundant st runt ural members, appropriately locatedbarriers such as splices, and selection of reduced st resses in critical areas inwhich slow rates of crack growt h must be maintained if such should occur.

REDUNDANT STRUCTURE

The double- dell a wing provides an excellent example of t he use ofredundant structure to at tain fail-safe objectives. The wing spars areclosely spaced and, on the tension side of the wing, the spar caps are inmultiple pieces so that damage 1 o one piece will not propagate into another.

STRUCTURA L DESIGN PHILOSOPHY 373

Should any spar cap and adjacent skin fail for any reason whatsoever, thesurrounding structure has sufficient reserve strength to carry the incre-mental loads. In addition, the multiple spars provide a very intimateconnection between the wing and the fuselage. Features of the wing de-scribed above are illustrated in Fig. 17.

Mounting of the engines is another area ill which fail-safe structuraldesign philosophy is applied. The supersonic transport engines will bemounted by a link system which is tolerant to the failure of any one link.Figure 18 shows schematically the concept of this engine-mounting syst em.Under normal conditions, the front left mount takes side, vertical, andthrust loads. The front right mount has resilient features and normallycarries no significant load in any direction. The aft top mount takes sideload only, and includes secondary side snubbers, normally unloaded. Bothaft side links take vertical loads only.

For failure situations, the front, right mount is capable of absorbing thefull limit loads in all directions with a limited motion of the engine. Theaft top mount secondary snubbers are designed to carry limit side loadswhile restricting side motion of the engine. Either of the aft side links issufficiently strong to carry the entire vertical load after failure of theopposite side link.

Various failures are provided for by this mounting system. For example,after a front, mount failure, the other front, mount takes over. If either aftside link should fail, the other side link carries vertical load and the resilientfront mount comes into action and takes torque loads. If the aft top mountfails, the secondary snubbers come into action and carry side loads. Inaddition to these fail-safe characteristics, the mounting system providesself-aligning features for ease of installation, and thermal growth of theengine is permitted without restraint. Although fail-safe, the mountingsystem is not, redundant and does not transmit airframe loads to the engine.

Windshields and windows are designed to stringent fail-safe require-ments. As illustrated ill Fig. 19, the windshield installation contains fourpanes. In addition, the windshield post is a multipiece construction sofailure from one piece cannot pmpagate to the Other. Protection for thepilot is provided for bird impact and strikes by four-inch hailstones at,speeds up to V(' at sea level. With any one single-load-bearing element ofthe windshield installation failed, it can still withstand maximum cabinpressurization. With two elements failed, safe flight of the airplane can bemaintained with a cabin pressure altitude of not more than 15,000 ft. Thecabin windows are also a four-pane construction as illustrated in Fig. 19.They are designed to meet the same fail-safe requirements as the wind-shield insofar as they are affected by cabin pressure. In addition, thewindows are designed with a snIall area (1i-in. diamel er, clear vision) to

374 FOURTHCONGRESS — AERONAUTICAL SCIENCES

balance with cabin air-source capabilities so as to limit- maximum cabin-pressure altit ude to a safe value even in the remote possibility of a completerupture of all elements of the window.

SURFACE DETAIL

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Figure 17. Wing construction.

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Figure 19. Windshield and cabin Windows.

BARRIERS

The appropriate use of splices provides an excellent inhibitor of cracksor damage which may occur in the st ructure. As an example, wing skinsare panehzed at, each of the main spars. Splices in critical areas of the fuse-lage will be provided to act as barriers t.o crack or damage propagation.The integrally stiffened structure of the tin is made from panels appropri-ately spaced so damage front one cannot, propagate across t he splice lineinto t he adjacent panel.


REDUCED STRESSES FOR SLOW CRACK PROPAGATION

In extensive portions of the structure, fail-safe objectives are achievedthrough the appropriate selection of stress levels to maintain slow rates ofcrack growth. An excellent example of this type of structure is the skin ofthe fuselage. Pressurization stress in the duplex annealed 8-1-1 titaniumskin is limited to 25,000 psi, not only to provide excellent fatigue charac-teristics but also to provide for a very slow rate of crack growth shouldsuch occur in the structure.

To substantiate the selection of this stress level, a fuselage panel wasconstructed approximately 5 ft by 10 ft of 8-1-1 titanium duplex annealed0.030 skin material with 0.050 rings and 0.032 stringers representative ofthe supersonic transport construction. This panel is shown in Fig. 20. Thecurved panel, mounted on a flat test bed and shrouded for hot-air heatingon the external surface, was cyclically loaded by internal air pressure to apressure of 12.5 psi, stressing the skin primarily in the circumferentialdirect ion.

The test temperature ranged from 550-650°F. Initial damage was pro-duced intentionally by a sawcut approximately 4 in. long, centrally locatedin the skin between two rings and two stringers. A true fatigue crack wasproduced at the ends of the sawcut at room temperature before testing atelevated temperature. This procedure was repeated twice to accelerate thetest. After the crack in the skin had progressed substantially beyond thetwo adjacent rings, the redistribution of the load caused these two rings tofail, and the crack growth rate increased. However, after the crack hadextended over three bays (crack length = 33.4 in., Fig. 20), the crackgrowth was arrested by the next pair of rings. The panel was then failedstatically at room temperature. The test showed that, even with this extentof damage, the design configuration would withstand an internal pressureof 12.25 psi.

From the recorded test data, an estimate of the number of pressure cyclesrequired to fail t he titanium fuselage structure from any initial crack lengthcan be made. For example, a 1-in , initial crack would require approximately46,000 cycles, or flights, to reach unstable proportions; or a 10-in , initial

crack would require approximately 2,100 cycles, or flights, to reach un-st able proportions. The 46,000 flights to failure from a one-inch crack areof the order of the required lifetime of the supersonic transport, while the2,100 cycles to failure from a 10-in. crack would approximate a year'sflying. This test provides substantiation of the excellent fail-safe charac-teristics of titanium 8-1-1 duplex annealed material for the supersonictransport environInent, and demonstrates the feasibility of design to lowcrack growth rates, thus insuring catching of such faults during routineinspections.


Figure 20. Fuselage crack investigation panel.

Figure 21 further illustrates the excellent crack resistance of titanium8-1-1 duplex annealed when compared with aluminum and stainless steel.The rate of crack growth in stiffened structure of various materials iscompared on a strength/density basis in which X is the crack length. At550°F, comparable structure of titanium 8-1-1 duplex annealed has anappreciably slower rate of crack growth than the aluminum alloys used incurrent subsonic transports.

FAIL-SAFE TESTING AND ANALYSIS

The attainment of actual fail-safe structure utilizing the three basictechniques described above is implemented by a combined program ofanalysis and testing. The allowable strengths of damaged structures aredetermined experimentally for the specific structural configurations andexternal environment of the supersonic transport. Simple and successfulempirical methods have been developed for translating experimental datainto useful and accurate analytical methods for aluminum structures.These design procedures have been proved in the design, testing, and servicerecord of current subsonic transports. Test programs are under way toprovide the necessary experimental data for the materials and environmentof the supersonic transport which are required to apply these analyticalprocedures. In addition, extensive tests of specific components will be


performed to further validate the analysis. These tests include windows,Windshields, components of the wing, and fuselage. These tests and analysisare the prelude to the full-scale t est s described in a later section.

INSPECTION AND MAINTENANCE

The fourt h fundamental ingredient of the supersonic transport structuraldesign philosophy is based upon positive recognition of the fact thatinspection and maintenance of st met ure must be considered from the out-set. As has been discussed previously, inspection and repair are consideredto be fundamental ingredients of both the fatigue and fail-safe aspects ofthe airplane. In this manner, inspect:ion and maintenance contribute towardt he achievement of the primary goal of unlimited service life. In additiont o these aspects, consideration of inspection and maintenance requirementsplays a large role in the selection of t he basic structural configuration ofmajor airframe elements and their detailed execution. For example, inspec-tion and maintenance requirements are responsible for the basic decisionto utilize essentially conventional, built-up aircraft st ructure for thesupersonic transport, rather than to make extensive use of Honeycomb orother sandwich-type structures.

6 X 105

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RATE OF CRACK GROWFM, IN./CYCLE

Figure 21. Rate of crack growth on stiffened structure.

sun-cm:HAL DESIGN PHILOSOPHY 379

The double-delta wing is well suited to the provision of adequate inspec-

tion capability. The extremely long root chord makes possible a structural

depth of the order of 5 ft in the root region, making it possible to provide

access to the interior of the wing where large cavities having all open-type

structure are available for simple visual inspection. The use of integral fuelin the inboard portion of t he wing in lieu of bladder or other fuel cells

facilitates visual inspection in this critical area. In addition, the multiplespar construction of the wing utilizes open trusses in many areas contrib-

uting to both accessibility and visual inspection. The built-up skin con-

struction is similar to that which has been used in previous transport

aircraft and can be repaired by conventional techniques. Leading andtrailing edge elements are both repairable and replaceable by detachment

from the primary-wing box structure.The fuselage is constructed in a conventional manner with a single skin

combined with rings and stringers. The single-skin thickness is inspectableon the exterior, and the stringers are of open section and available for

inspection upon removal of cabin trim and insulation bats.The adoption of these essentially conventional and well-proven ap-

proaches to transport aircraft structural inspection and maintenance insure

a maximum carryover of the knowledge and techniques developed fromyears of operation. The principal new element is the introduction of the

new materials, an area which has minimum effect on airline procedures,

experience, and cost.

STRUCTURAL PROOF APPROACH

The final element of the st ructural design philosophy for 1 he supersonictransport is a very comprehensive test program. Flight-loads measurement,

ultimate-strength tests, fail-safe strength tests, and fatigue tests will beconducted. The static and fatigue tests demonstrate the strength and life

of the airframe for the loads imposed. The test loads will be verified byextensive and accurate loads measurements in flight and ground operations

of t he aircraft..

FLIGHT-LOADS PHOGIZA

A flight-test program will be conducted on the first airplane with theobjective of demonstrating the structural integrity of the supersonic

t ransport. This program will include measurement of external applied loads

and temperatures by flight-load survey methods; measurement of dynamicloads for various conditions such as rough air, landing, and taxi; and


measurement of internal load distributions and detail stresses for bothexternal applied loads and loads induced by structural heating. This pro-gram will culminate in demonstration-type maneuvers to limit conditionsfor selected maneuvering conditions. Airload survey instrumentation willinclude provisions for obtaining fuselage and wing loads with both pressuredistribution measurements and calibrated load-measuring strain gageinstruinent at ion.

All load measurements will be available to correlate measured loads withdesign loads before conducting the ultimate load static tests. Any adjust-ments in loads or load distributions required to make the static test moreproperly reflect the actual flight loads can be made prior to completion ofthe static tests. The airplane fatigue-test program starts late enough in thedevelopment program to allow full advantage to be taken of the structuralflight test results.

A flight-flutter test program will be conducted with the objective ofdemonstrating freedom from flutter for any flight condition and airplaneloading at speeds up to VD/u1/D.In addition, a substantial amount of datawill be obtained to compare with design flutter data to enable verificationthat satisfactory flutter safety margins are obtained.

ULTIMATE STRENGTH TESTS

Ultimate strength tests and fail-safe demonstration t.ests are to beconducted on one airframe. These static tests will be conducted in thefatigue-loading apparatus. The reason for this is that the fatigue-testequipment has evolved into a very flexible, easy changeover system. Mag-netic tape controlled hydraulic servo valves control the load distributionover the airframe. Thus, changing to a different loading condition is nolonger the time-consuming project of shifting jacks, repiping, and recheck-out. A new preprogrammed load-control magnetic tape is sufficient toaccomplish the complete change to a new loading case. Testing moreconditions with a considerable savings of time and expense is realized bythis approach.

Heat. will be introduced, where required, by program-controlled thermo-couple monitored quartz lamps. Cooling water-fog and forced air distribu-tion will remove heat at t he required rate. Automatic lockup circuits areprovided to save 1 he test article in case of malfunction overload or failure

in any portion of t he test equipment or airframe. Automatic data readoutand recording systems will be used for compatible loading and heatingtimes. Selected channels will be monitored throughout the test.


FAIL-SAFE TESTS

Fail-safe tests will be conducted on the static test airframe on completionof the ultimate-strength series described above. These tests will be con-ducted both statically, with intentional cuts introduced prior to loading,and dynamically, with the cuts introduced suddenly while the structure isloaded. Heat, pressurization, and static loading will be applied as necessaryto represent the requirements of the specific condition.

FATIGUE TESTS

The fatigue philosophy described previously relies heavily on accuratelaboratory simulation of service loading and environmental conditions, andto a lesser degree on theoretical fatigue analyses. As a result of this emphasis,a refilled fatigue test system has been evolved. Magnetic-tape-controlledhydraulic servo valves with load cell or position feedback loops have provedvery practical and versatile. Flight-by-flight loading sequences are used toavoid arbitrary decisions for the definitions of the very sensitive ground-air-ground cycle per flight.

Figure 22. P-3A wing landing gear fatigue tests.


This technique has been developed extensively and used to complete afatigue t est program for the United States Navy on the P-3A antisubmarinepatrol airplane. A photograph of the test setup appears in Fig. 22 wherethe tape-controlled loading jacks are clearly evident. Over 2,000 flights,complete from takeoff to landing with such detail operations as flap exten-sion and retraction, propeller loads, and landing-gear loads, were completedin a calendar period of approximately 4 months. This technique is fullydeveloped and ready for application on the supersonic transport except forthe additional need to simulate thermal cycles.

In addition to the influence of heat on the material properties, heatingand cooling of the structure during supersonic flight is an important sourceof fatigue stress cycles. Ideally, since parts of the structure do not reachtemperature equilibrium during service flights, each fatigue-test flightshould be carried out on a real-time basis. Practically, this is not possiblein a test program which must provide proof of adequacy within a limitedtime. Long-time experimental programs are now underway to determinethe possible acceleration of the heating cycles during fatigue tests withoutundue compromise on the results of the test.

CONCLUSION

This paper has presented a broad view of the structural design philosophybeing employed on the supersonic transport. As indicated at the outset,this philosophy is based upon the five elements of rational load criteria,unlimited life, fail-safe for both strength and stiffness, inspectability andmaintainability, and complete structural proof testing. In each of theseareas, the approach adopted is to carry over the extensive knowledge andexperience which has been gained over the years and to apply it in the newenvironment and to the new materials as required. By taking full advantageof experience and full account of new conditions, the detailed execution ofthis philosophy will result in a supersonic transport of sound structuralcharacteristics. The depth of knowledge available from experience and thepromise shown by the new materials leads to the conclusion that thesupersonic transport will have even greater reliability and safety than thatassociated with the current subsonic transports.

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STRUCTURAL DESIGN PHILOSOPHY - ICAS 353 Heppe et al.pdfExisting discret e gust criteria are known to provide satisfactory structure for aircraft similar in configuration, flight profile,