NASA Technical Memorandum 4768

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Titanium Honeycomb

Panel Testing

W. Lance Richards and Randolph C. Thompson

October 1996

NASA Technical Memorandum 4768

Titanium Honeycomb

Panel Testing

W. Lance Richards

Dryden Flight Research Center

Edwards, California

Randolph C. Thompson

PRC Inc.

Edwards, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1996

TITANIUM HONEYCOMB PANEL TESTING

W. Lance Richards

NASA Dryden Flight Research Facility

Edwards, California

Randolph C. Thompson

PRC Inc.

Edwards, California

ABSTRACT

Thermal-mechanical tests were performed on a titanium

honeycomb sandwich panel to experimentally validate the

hypersonic wing panel concept and compare test data with

analysis. Details of the test article, test fixture development,

instrumentation, and test results are presented. After extensive testing to 900 *F, nondestructive evaluation of the

panel has not detected any significant structural degradation

caused by the applied thermal-mechanical loads.

NOMENCLATURE

A1 aluminum

DACS Data Acquisition and Control System

LID liquid interface diffusion

Mo molybdenum

NDI nondestructive inspection

Si silicon

Sn tin

Ti titanium

TiHC titanium honeycomb core

V vanadium

Zr zirconium

A T temperature change from initial reference

temperature, *F

INTRODUCTION

Honeycomb-core sandwich panels were identified in the

late 1960's as one of several candidate concepts for use

on high-speed aircraft [1]. In early investigations, honeycomb panels exhibited problems with the bonding between

the facesheets and the honeycomb core. Improved bonding

techniques, such as the liquid interface diffusion (LID) process (LID bonding is a Rohr Proprietary process), have resuited in the reemergence of titanium honeycomb panels as

leading candidates for the wing panels of a Mach 5 aircraft.

Personnel at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Facility (DFRF)

have recently completed thermal-mechanical tests on two

LID-bonded titanium honeycomb panels. The test program

was based on a study conducted at the NASA Langley Research Center (LaRC) in which two titanium wing structure concepts were analyzed for use on a Mach 5 vehicle

[2]. This study led to the development and fabrication of

two LID-bonded titanium honeycomb wing panels designed

to survive the thermal stresses produced in a 900 *F thermal environment. A test program was developed at NASA

DFRF to demonstrate the wing panel design and fabrication,

evaluate the LID bonds between the core and facesheets, and

correlate test data with a finite-element analysis performed

at NASA LaRC. The test and modeling techniques were

refined before exposing the panel to the ultimate 900 *F

thermal environment. This ensured that the thermal stresses

produced in the panels were representative of Mach 5 flight

conditions and that preliminary test data and analysis were

compatible. This paper describes some of the difficulties

overcome in the testing and analysis to successfully demonstrate these wing panels for use on a Mach 5 vehicle. Details of the panel evaluation tests including the test article, instrumentation, test setup development, and results

are discussed.

TEST OBJECTIVES AND

METHODOLOGY

The titanium honeycomb-core ('EHC) wing panels were

required to survive the thermal stresses produced in Mach 5

flight. Figure 1 depicts such a wing panel as part of a multipanel array in the outboard wing section of the airplane.

During the aerodynamic heating encountered in a typical

mission, each panel will be prevented from rotating about

its edges by the surrounding wing structure. To simulate

these flight conditions in the laboratory, the preferred testing

approach is to mount the test article with additional buffer

panels on a fixture representing the wing structure. The test

article, placed in the center of the multipanel array, is less

sensitive to unrealistic boundary effects in this arrangement.

When simulated flight temperatures are applied to the entire
structure, realistic thermal stresses develop naturally in the

center panel. Although this approach is highly desirable and

has produced successful results [3], it is also very expensive.

An alternative test approach was pursued in this program.

Because of budget constraints, two "HHC wing panels were

tested separately and not in a multipanel array. The same

ideal thermal-structural loads determined in Ref. [2] were

applied in a single panel test setup. The original test goals

were to apply unifcxa-n temperatures across the upper wing

panel skin, prevent all four panel edges from rotating, and

allow the panel to thermally expand in-plane. It was recognized from the outset that producing the proper boundary

conditions on a single panel would be very difficult, if not

impossible to achieve.

The test program was divided into mid- and hightemperature thermal-mechanical tests. The first phase was

to thermally cycle the wing panels 50 times from room temperature to 600 *F. The second phase was to thermally cycle

the wing panels 50 times from room temperature to 900 *F.

Exposing the panels to the 100 transient thermal-stress cycles also was done to assess low cycle fatigue performance.

Nondestructive inspection (NDD was performed after manufacturing and after each of the 50 thermal cycles to verify

the core to facesheet bonding integrity.

TEST DESCRIPTION

Test Article

Figure 2 shows one of two "HHC sandwich panels

built under a contract to Lockheed Aeronautical Systems Co. (Calabasas, California) and Rohr Industries, Inc.

(Chula Vista, California). The panels measure 23 in.

square and consist of two 0.060-in.-thick titanium alloy

(q_-6A1-2Sn-4Zr-2Mo-.09Si) ('I1-6242-Si) facesheets LIDbonded to a 0.69-in.-thick honeycomb core. The core is

formed into 0.1875-in. hexagonal cells from 0.002-in.-thick

titanium alloy (Ti-6A1-4V) foil. Edge closeouts are brazed

to the core on all four edges and are made from 0.020-

in.-thick Ti-6242-Si. Ti-6AI-4V bearing plates measuring 1.5 in. square and 0.125 in. thick were fastened to the

perimeter of the upper and lower panel surfaces.

Instrumentation

The panels were instrumented with sensors capable of

measuring surface strains, temperatures, and out-of-plane

deflections at 600 and 900 *F. The instrumentation included conventional foil strain gages arranged in a rectangular rosette for use at 600 °F. Foil strain gages and weldable

strain gages were used at 900 *F. Duplex glass braid insulated type K thermocouples were used at 600 and 900 *F.

Deflection potentiometers also were used throughout the test

program to measure out-of-plane panel deformations. The

sensors were installed on one quadrant of the upper and

lower facesheets to utilize the two planes of symmetry that

exist about the panel centerline axes. Figure 3 shows the

upper surface of a panel instrumented for testing at 900 *E

Data Acquisition and Thermal Control

Data acquisition and adaptive digital thermal control were

accomplished by using the Data Acquisition and Control

System (DACS) Thermostruetures Research Facility at the

NASA DFRF [4]. The DACS was used to apply the same

600 and 900 *F temperature profiles to the upper surface

of the panel as determined in a heat transfer analysis [2].

A thermocouple located at the center of the upper surface

was used in the DACS feedback algorithm to impose a predicted flight temperature profile by varying the power supplied to the heaters. For preliminary testing, thermocouples

located on the side closeouts were used to control the edge

temperatures.

The DACS maximum allowable system measurement error is 4-0.15 percent of reading or 4-20 _V, whichever is

greater. Therefore, for a 4-20 #V strain measurement input

from a single active arm strain gage with a 4-V direct current

(DC) excitation voltage, the error band is 4-8 _in/in. However, this error is reduced with additional active arms and

higher excitation voltages. Similarly, a type K thermocoupie measurement error with a 4-20 _V input is equivalent to

+0.9 *F.

TEST DEVELOPMENT

This section describes the process used to develop a test

to adequately achieve the program objectives.

Initial Test Setup Description

Figure 4(a) depicts the in-flight configuration of a titanium wing panel mounted on a wing structure. Figure 4(b)

shows the single-panel concept used in this test. The top and

center schematics show that when one surface of a simplysupported wing panel is heated, the panel will bend out-ofplane to alleviate thermal stresses. The bottom schematic

shows the boundary conditions desired in the test. The goal

was to prevent rotation at the panel edges while allowing inplane translation. The upper and lower bolts shown in Fig. 5

were offset to react against panel rotation as the upper surface of the panel was heated. The 0.75-in. perpendicular

distance between the bolts was designed to maximize the

moment restraint and minimize thermal shading at the panel

edges from the overhead heat source. Fifty-two separate restraint mechanisms were located around the perimeter of the

panel (13 on each side). The restraint mechanisms consisted

of a 0.5-in. by 0.75-in. by 1.375-in. stainless steel load pad

and UNF 3/8-24 high-strength steel bolts. The end of each

bolt had a spherical radius that fit into a spherical socket in

the load pad. The load pads, bearing plates, and bolts are

shown in Fig. 5. The bolts in turn were supported by large

structural steel members consisting of angle beams (L6 X

6 X 0.75), two 0.75-in.-thick back plates and two to three

C beam stiffeners. Angle beam stiffeners were welded to the
anglebetweeneverythirdor fourth bolt. High-temperature

lubricant, located between the load pads and bearing plates,

allowed the panel to thermally expand in-plane. Small goldplated reflectors were used to heat the sides of the panel and

help offset the conduction losses at the panel edges caused

by the massive test fixture.

The upper facesheet surface was heated by infrared quartz

lamp heaters located 6 in. above the test article. Figure 6

shows the array of heating units used. Eight gold-plated reflectors, each containing 6 12-in.-Iong lamps, were capable

of producing 48 kW of total power.

Comparison of Preliminary Test Data With

Finite-Element Analysis

Both panels were exposed to 50 transient heating cycles

from room temperature to 600 *F in the initial test setup.

Temperature at the center of the upper surface was controlled to within 4-5 *F of the 600 *F transient heating profile calculated in Ref. [2]. The side heaters were also used to

heat the panel edges in an attempt to reduce heat loss at the

boundaries. The edge restraint mechanisms were brought

into contact with the panel edges to react against the panel

edge rotation as the panel thermally distorted during the

heating profile.

After the initial testing, the preliminary data were compared with a finite-element analysis performed at NASA

LaRC. Preliminary data-analysis correlation showed that

the analytical strains were significantly higher than the measured strains. The analysis also predicted that the magnitudes of the compressive upper surface strains were equal

to the lower surface tensile strains. However, the measured

upper surface strains were almost twice the magnitude of the

lower surface strains. The measured and calculated deflections at the panel center disagreed by 16 percent [5].

Evaluation of Preliminary Test Techniques

The following testing issues required further investigation to understand why experimental and analytical results

did not compare well.

Strain Measurements. One of the largest uncertainties

in the strain gage measurement during the 600 OFtesting was

the apparent strain in the gage installation. Apparent strain

is defined as the strain produced by the difference in thermal

expansion between the gage and the material to which it is

attached. A smaller component of apparent strain is caused

by the change in gage factor and the temperature coefficient

of resistivity of the gage material with temperature. Norreally, apparent strain is characterized by installing strain

gages on coupons in which both the gage and the coupon

are assumed to represent those used in the test. After the

instrumented coupon is heated several times to the expected

temperature range, the apparent strain correction curve is

obtained by averaging the thermal cycle data sets. Figure 7

shows apparent strain as a function of temperature for a

bonded foil strain gage on Ti-6242-Si. This curve is then

used after testing to correct the test data for apparent strain

at any given test temperature.

Apparent strain tests were originally performed on

coupons made from the same material as the panel

facesheets. To verify that the strain gages were wellcharacterized for apparent strain, additional apparent strain

tests were conducted using the instrumented test article instead of coupons. The advantage in this method was that

the actual strain gages and the test article could be uniquely

characterized for apparent strain. This approach avoids assumptions about how well the coupon tests represent the

gages and test article used in the heating tests. The apparent strain correction technique assumes, however, that the

stress-induced strain in the built-up panel is negligible and

that only the apparent strain is measured. The maximum difference between the panel and coupon apparent strain curves

was approximately 90 vin/in. The data precision was also

improved from 4-50 to 4-25 t_in/in, by repeating the apparent strain tests.

Boundary Conditions. Another uncertainty in the test

was the degree of end fixity the test fixture provided in the

prevention of the panel edge rotation. Although the test fixture was to simulate a perfectly fixed end condition, achieving perfect end fixity in the laboratory is almost impossible

[6]. This is especially true when thermal loads are applied

because of the thermoelastic nature of materials. The true

edge restraint provided by the test fixture was necessary to

produce an accurate finite-element model of the panel.

Several tests were conducted to resolve the unknown test

fixture stiffness. Figure 8 shows the test setup to quantify

the degree of end fixity provided by the test fixture. An

aluminum beam was installed in the test fixture and subjected to a mechanical load by the hydraulic jack shown at

the bottom of the figure. The beam deflection was measured by dial gauges shown above the beam. The load

cell shown at the bottom of the figure was used to measure

the mechanical load applied to the beam. The same moment expected in the heated panel was used to determine the

equivalent mechanical load to be applied to the aluminum

beam. Comparisons of the measured deflections with simple beam theory showed the fixture provided approximately

35 percent of a fixed beam. These tests explained why the

measured strains in the wing panel tests were much lower

than those predicted from an analysis which assumed perfeet end fixity.

Temperature Distributions. Additional heating tests

were required to isolate how the top and side heaters individually affected the strain gage measurements. In these

tests, the top heating profile was applied to the panel without

heating the panel edges. Likewise, the side heating profile

was applied to the panel edges without heating the upper

panel surface.

The top heating test revealed that in-plane thermal

gradients caused by the shaded panel edges produced
compressive stresses to the panel. These compressive

stresses added to the upper surface compressive stresses

and subtracted from the lower surface tensile stresses. This

explained why the magnitude of the upper surface strains

was larger than the magnitude of the lower surface strains.

Test results also showed that although the side heaters produced more uniform temperatures on the upper panel surface they also reduced the magnitude of the stresses in the

upper facesheet.

Evaluation-Modification of Preliminary Modeling

Techniques

The test setup design process also uncovered analytical

issues that led to the refinement of the finite-element model.

These refinements, originally described in Ref. [5], included

the following improvements. (1) The model was refined by

utilizing the two planes of symmetry in the panel. Although

the model was reduced from a full-panel model to a onequarter-panel model, the number of degrees of freedom in

the analysis were increased tenfold. (2) The axial rod and

shear web elements used in the original model were replaced

with plate and solid elements. The plate elements were used

to represent the panel facesheets and the solid elements represented the honeycomb core. These elements enabled the

use of anisotropic stiffness coefficients to better model the

behavior of the honeycomb core. (3) Another significant

improvement to the test data and analysis correlation was

gained when the actual test temperatures were used in the

model instead of the constant temperature distribution assumed in the original model. Therefore the refined analysis

was able to more appropriately simulate the stresses produced by the large in-plane temperature gradients measured

in the test. (4) The new finite-element model also included

temperature-dependent physical properties which were not

used in the initial model.

Final Comparison of Preliminary Test Data With

Refined Finite-Element Analysis

Although the panel stresses were far lower than the stress

levels predicted in Ref. [2], major improvements were made

that resulted in better correlation between the experimental

data and the analysis. This was essential before conducting the 900 *F evaluation tests. The measured upper surface strains were in good agreement with the calculated values. The maximum difference between measured and calculated lower surface strains was 65 vin/in. The measured

and calculated center panel deflections differed by less than

5 percent.

Final Test Setup Description

The aluminum beam test results showed that the test fixture used in the original 600"F thermal cycle tests was not of

sufficient stiffness to produce the stress levels necessary to

validate the wing panel concept. The test fixture was modified to improve this stiffness. Figure 9 shows the stiffened

test fixture moment restraint. The most significant modi4

fication to the fixture was the removal of the side heaters

because this allowed the welding of a 0.75-in.-thick steel

plate close to the panel edges. Consequently, the moment

arm was reduced by approximately 75 percent, greatly improving the test fixture stiffness at the panel attachment area.

Figure 10 compares the initial and final test fixture configurations. Seventy-two additional angle beam stiffeners were

welded to the fixture to prevent local deflections in the angle beams at the restraint mechanisms. Large structural steel

I beams (W8 X 24) were welded to the back plates to further

increase the test fixture stiffness.

Figure 11 presents the edge-moment calibration test resuits for the final test fixture. The center deflection of a

simply-supported and a fixed-beam under the application

of a centrally-located concentrated load were determined

from simple beam theory as functions of load. The simplysupported and fixed-beam deflection curves are depicted by

the uppermost and lowermost lines, respectively. The experimental data from the beam tests are shown between these

two theoretical curves. This figure shows the improvement

in test fixture stiffness achieved with the final test fixture.

The test fixture edge fixity was improved from 35 to 75 percent of a perfectly fixed restraint.

Test Procedure Modifications

The test procedures were also modified as a result of the

test technique evaluation studies. Instead of trying to impose the anticipated in-flight heating and boundary conditions to the panel, the test procedures were changed to impose the stress levels predicted in Ref. [2]. These stresses

were produced in the panels by increasing the panel heating rate and applying additional bending moments at the

panel edges. The combination of mechanical and thermal

stress was varied until the proper stresses were produced in

the panel. The restraint mechanisms, originally designed to

simply react against thermal distortion, were used to mechanically induce a moment at the panel edges. The heating rate imposed in the preliminary tests was increased from

3.5 to 15 *F/see to produce a higher thermal gradient and

hence higher compressive stresses in the facesheet.

RESULTS AND DISCUSSION

The final test fixture was designed and fabricated to provide increased edge fixity. Test procedures were modified

and several 600 °F checkout tests were performed before 50

of the 900 *F cycles were run.

Figure 12 shows the measured strain time history at the

center of the panel on the upper and lower facesheets for

a representative thermal cycle to 600 *F. Before heat was

applied to the panel, a preload was applied to the bolts

to induce a mechanical moment on the panel. The strain

gage output resulting from this mechanically-induced load

is shown at the beginning of the profile in Fig. 12. At approximately 100 sec, a 15 °F/see temperature profile was applied to the upper surface of the panel. The heating profile
appliedtotheupperfacesheet causedatemperaturegradient

(AT) through-the-thickness and in-the-plane of the panel.

A schematic of the heating condition and the resulting temperature distribution from a representative test is shown in

Fig. 13. The panel temperature sharply declines at the edges

because the test fixture shades the panel border from radiant heat. The test evaluation studies showed that these

in-plane thermal gradients cause significant compressive

stresses which add to the upper surface compressive strains

and subtract from the lower surface tensile strains. The upper surface strains shown in Fig. 12 are larger in magnitude

than the lower surface strains because of the additional compressive stresses produced by the in-plane A T.

The measured strains were converted to stress values using a strain gage rosette reduction analysis. Table 1 compares these experimentally-determined stresses with the target stresses necessary to validate the concept. This table

shows that the upper facesheet stress levels achieved in the

final test setup were in 2-percent agreement with the target

stress. The lower facesheet stresses do not compare as well

because of the compressive stress produced by the in-plane

A T. However, the upper surface center stresses produced

in the final test fixture moment restraint were sufficient to

demonstrate the panel concept.

After the 600 °F checkout tests were conducted using the

final test fixture, 50 of the 900 *F thermal cycles were conducted on the same panel. The experimental data and analysis correlation from these tests is ongoing. This panel was

subjected to NDI before and after the 900 °F thermal cycles.

Figure 14 is a photo of an x ray of one quadrant of the panel

after 50 of the 900 *F cycles. Examination of the test results

has not revealed any core-to-facesheet disbonding.

CONCLUDING REMARKS

A titanium honeycomb-core sandwich panel has been

thermally and mechanically tested to evaluate the panel concept and compare the results with analysis. The test techniques described were used to test the panel to 600 and

900 °F at predetermined stress levels. The panel survived

more than 100 of the 600 °F thermal cycles and 50 of the

900 *F thermal cycles. Heating profiles were applied to the

panel by radiant heat using quartz lamps and gold-plated

reflectors. Initially, the measured strains were lower than

finite-element analysis because of difficulties in both simulating and modeling complex boundary conditions. The

test and modeling techniques were refined to ensure that the

thermal stresses produced in the panels were representative

ofMach 5 flight conditions and that preliminary test data and

analysis were compatible. Stress levels required to validate

the wing panel were obtained by varying the mechanicallyinduced bending moments at the panel edges and by varying

the through-the-thickness temperature gradients. In-plane

thermal gradients on the upper facesheet increased the upper

facesheet compressive stresses to the stress levels necessary

to achieve the test objectives.

REFERENCES

1. Plank, P.P., Sakata, I.F., Davis, G.W., and Richie,

C.C., Hypersonic Cruise Vehicle Wing Structure Evaluation,

NASA CR-1568, 1970.

2. Taylor, Allan H., Jackson, L. Robert, Cerro, Jeffrey A.,

and Scotti, Stephen J., "Analytical Comparison of Two

Wing Structures for Mach 5 Cruise Airplanes," J. of Aircraft, vol. 21, no. 4, Apr. 1984, pp. 272-277.

3. Fields, Roger A., Reardon, Lawrence F., and Siegel,

William H., Loading Tests of a Wing Structure for a Hypersonic Aircraft, NASA TP-1596, 1980.

4. Zamenzedah, Behzad, Trover, William F., and Anderson, Karl A., "DACS II - A Distributed Thermai/Mechanical

Loads Data Acquisition and Control System," Proceedings

of the International Telemetering Conference, San Diego,

CA, Oct. 26-29, 1987, 1987, pp. 737-752.

5. Jones, Stuart C., and Richards, W. Lance, "Titanium

Honeycomb Panel Thermostructural Test and Analysis,"

Workshop on Correlation of Hot Structures Test Data With

Analysis, NASA CP-3065-Vol. II, 1990, pp. 118-175.

6. Roark, Raymond J., and Young, Warren C., Formulas for

Stress and Strain, Fifth ed., McGraw-Hill Book Co., NY,

1975, pp. 325.

Table 1: Comparison of target and experimentally-determined stresses

Stress, lb/in 2

Upper surface Lower surface

Target -25544 25127

Experiment-final fixture -26100 18700
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Fig. 14: _HC panel x ray

17
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October 1996 Technical Memorandum

4.TITLE AND SUBTITLE i 5. FUNDING NUMBERS

Titanium Honeycomb Panel Testing

6. AUTHOR(S)

W. Lance Richards and Randolph C. Thompson

7.PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

WU 505-63-40

8. PERFORMING ORGANIZATION

REPORT NUMBER

H-2133

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM-4768

11. SUPPLEMENTARY NOTES

Presented at the Structural Testing Technology at High Temperature Conference, Society for Experimental Mechanics, Inc.,

Dayton, Ohio, Nov. 4--6, 1991. W. Lance Richards, NASA Dryden Flight Research Center, Edwards, California; Randolph

C. Thompson, PRC Inc., Edwards, California.

121. DISTRIBUTION/AVAILABILITY STATEMENT 12b. DI_III/BUTION CODE

Unclassified--Unlimited

Subject Category 39

13. ABSTRACT (Maximum 200 words)

Thermal-mechanical tests were performed on a titanium honeycomb sandwich panel to experimentally

validate the hypersonic wing panel concept and compare test data with analysis. Details of the test article, test

fixture development, instrumentation, and test results are presented. After extensive testing to 900 °F, nondestructive evaluation of the panel has not detected any significant structural degradation caused by the applied

thermal-mechanical loads.

14. SUBJECTTERMS

Elevated temperatures, Hypersonic vehicle, Strain gages, Structural analysis,

Structural testing, Test techniques

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