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~ PEN00260 J REPORT MDCA3265

NASA FLIGHT EXPERIMENTS

X-24C

FINAL BRIEFING

-}-.--_______________ -----l

... 1

N1CDONNELL AIRCRAFT COIWPANY .. /

MCDONNELL DOUG~ CORPORII&TION

J

REPORT MOC A3265 20 JANUARY 1975

NASA FLIGHT EXPERIMENTS X-24C

FINAL IBRIEFING

Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b), 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

COPY NO. ----!/-'.Z'-_

/MCDONNELL AIRCRAFT CO/MPANY

Box 516, Saint Louis, Missouri 63166 - Tel_ (314)232-0232

/

_CDONNELL DOUG~ CORPORATION

------~-

REPORT Moe A3265 20 JANUARY 1975

FOREWORD

This report summarizes the results of an investigation performed

at McDonnell Aircraft Company (MCAIR), a division of McDonnell Douglas

Corporation, P.O. Box 516, St. Louis, Missouri, 63166, to define and

cost an actively cooled fail safe structural system flight experiment

and an integral liquid hydrogen fuel tank structural system flight

experiment for the X-24C research aircraft program. The study was

conducted under NASA Contract NASl-13631 with a period of performance

from 8 November 1974 to 20 January 1975.

The principal investigators were Mr. A. H. Baker on the hydro

gen tank experiment and Mr. R. L. Herring on the active cooling ex

periment. This activity was conducted within the MCAIR Hypersonic

Aircraft Systems group managed by Mr. C. J. Pirrello. The Hypersonic

Aircraft Systems group is part of the HCAIR Advanced Aircraft Engi

neering Division directed by Hr. H. D. A1tis. Cost estimates were

conducted within the Contracts Division under the direction of

Hr. D. T. Mueller.

LIST OF PAGES

Title Page 1 through 175

MCDONNELL AIRCRAFT COMPANY

:1>ZZ oi UlO Co $:c $:(") :l>j :co -<Z

REPORT Moe A3265 20JANUARY 1975

SCOPE

Each experiment was investigated separately. For each, a

conceptual design of the airframe structure was developed along

with a conceptual design of the associated functional subsystem

elements. The design, development and testing required to assure

the flightworthiness of each experiment was determined, with the

ground rule that the experiment be capable of field installation

at the test site in the basic X-24C research aircraft. The manu

facturing plan was based on an experimental shop approach. Flight

test requirements were defined and a flight test plan developed.

Where appropriate critical technologies and long lead development

on procurement items were identified. Finally, the overall pro

gram schedule was developed and rough order of magnitude costs

were determined in constant 1974 dollars.

IKCDONNELL AIRCRAFT COIKPANY

2

REPORT MDe A3265 20 JANUARY 1975

SCOPE

• DEVELOP A CONCEPTUAL DESIGN

• DETERMINE DESIGN, DEVEL REQUI REMENTS

.• DEFINE MANUFA

~.IlP?\L TECHNOLOGIES AND LONG

• PROVIDE R.O.M. COSTS AND SCHEDULE GP74-1039-24

MCDONNELL AIRCRAFT COMPANV

3


STUDY APPROACH AND ASSUMPTIONS

Due to the short study period involved, the conceptual design

was developed with a minimum of analytical effort. Good engineer

ing judgement was used based on the results of related studies

currently in progress at MCAIR. It was assumed that all the detail

design and performance information on the basic X-24C airplane

would be readily available.

Structural and functional interfaces were assumed to be estab

lished at go-ahead and coordinated by the X-24C vehicle contractor.

All provisions for the experiments were assumed to be previously

incorporated in the basic X-24C, thus no costs were included to

cover modifications to the basic vehicle to accommodate the exper

iment.

MCDONNE ...... AIRCRAFT COMPANY

4


STUDY APPROACH AND ASSUMPTIONS

• CONCEPTUAL DESIGN

• NO STRESS OR WEIGHTS ANALYSIS

• MINIMAL THERMAL AND SYSTEM ANALYSIS

• NO TRADE STUDIES - USE JUDGMENT

• BASIC X-24C DETAIL AIRFRAME AND SUBSYSTEM DESIGN AND PERFORMANCE REQUIREMENTS ARE AVAILABLE

• INTERFACES ESTABLISHED AT ATP

• STRUCTURAL PROVISIONS FOR EXPERIMENTS INCORPORATED IN

BASIC X-24C

• SYSTEM CONTRACTOR COORDINATES INTERFACES GP74-1039-12


5


EXPERIMENTS GENERAL ARRANGEMENT

The integral liquid hydrogen tank is an all-welded design of 2219 aluminum. The

shell of the multi-bubble design is integrally stiffened in an "isogrid" pattern. The

intersecting cones forming the bubbles provide straight-line load carrying elements

and have planar webs at each intersection to carry shear and react internal pressure

loadings. Elliptically domed tank ends, one of which includes an access door, are in

cluded to increase volumetric efficiency at minimum weight. Thermally induced changes

in tank size with respect to basic aircraft structure are accommodated by motion of the

monoball-ended support l~nks. Load carrying continuity is not affected by this motion.

Operational elements of the liquid hydrogen system such as the supply port, vent

and dump lines are identified here. Provisions for these lines must be made in the

basic x-24c aircraft. The helium bottle located aft of the tank is the source of pres

surant and purge gas necessary to function of the system.

The actively cooled fin structure is designed to operate at an average surface

temperature of 250°F, representative of actively cooled, hypersonic cruise aircraft

concepts. The fin leading edge selected is a low drag cooled concept. The active

cooling system is configured with an efficient coolant utilizing hydrogen as the heat

sink. The integrated system includes a failure detection system, which will provide

pilot warning"of malfunction. Provisions must be made for routing of coolant lines

between the fin and payload bay, the failure detection system (FDS) control unit, en

vironmental control of the control unit, wiring between the fin and payload bay and

between the payload bay and cockpit. Aircraft power must be supplied to the FDS and

the cooling system.

6


EXPERIMENTS GENERAL "ARRANGEMENT

ACTIVELY COOLED FIN

INTEGRAL LH2 TANK"

~-

GP74-1039-106


7


THE VALUE OF FLIGHT TESTING

The ques tion has often been asked, "Why mus t we flight test, \vhy can't we learn all

\ve need to know from ground testing?" Decision makers would like a simple straightforward

answer to a very involved question. In our view, the three Cs, Commitment, Confirmation

and Confidence, are answers to the question. In the scientific technological community,

knowledge is only fact after many observations and experimental verifications have been

achieved. Knowledge is not accepted on the basis of faith. It has been said that "one

test is worth a thousand expert opinions."

It is evident that the value of a flight test program is derived not only from the

laboratory development and testing which precedes the flight tests. However, the flight

test program demands a level of commitment which will yield flightworthy hardware. not

just a lab specimen, and focuses a necessary level of physical and financial resources to

the program.

Neither ground test nor analyses can predict the response of an aircraft in flight.

Thus, a value of flight test is Confirmation. The response of the aircraft will be to

the real environment, facts not judgement will be available. If there are any major de

sign flaws they will be uncovered and solved in a cost effective manner.

Scaling effects and correlation of predictions to actuals will contribute in a great

measure to the most precious commodity of confidence. Decision makers require reasonable

assurance of success and budgetary responsibility before committing to a major program.

Certainly a level of confidence is available from a strictly theoretical analysis of a

particular functional system. However, Confidence is increased when experimental labora

tory knowledge is added to the data base. But only through the repeated successful demon

stration of the functional performance of the system will the unknown then become the

design practice of production systems.


8


THE VALUE OF FLIGHT TESTING

• COMMITMENT • HARDWARE - NOT PAPER OR LAB EXPERIMENTS

• FOCUS RESOURCES

• CONFIRMATION • RESPONSE OF AIRCRAFT IN REAL ENVIRONMENT

• UNCOVER AND SOLVE A MAJOR DESIGN FLAW

• FACTSNOTJUDGMENT

• CONFIDENCE • FROM UNKNOWN TO DESIGN PRACTICE GP74-1039-10

• NO SURPRISES MCDONNELL AIRCRAFT cOItifPANY

9


STRUCTURAL FLIGHT EXPERIMENTS

Accepting the need for flight testing of advanced structural tech

nologies we are faced with an equally perplexing question, "Hhat should

be the nature of a structural experiment?"

Certainly no one will argue with the premise that it should be flight

weight not boiler plate structure. In our view there is a direct corre

lation between size of the structural specimen and the confidence gained

in the technology being demonstrated--small gains with small structural

panels and large gains with large structural assemblies. A major

structural assembly will simulate all its intended design functions.

It can house the required furnishings and functional systems, provide a

suitable aerodynamic surface, sustain loads, heating, vibration and

acoustics and provide the required stiffness. A major structural assembly

will be exposed to many of the handling and maintenance operations asso

ciated with operational systems, whereas small panels are usually treated

with kid gloves. Thus, in our view, small panels are suitable to perform

exploratory research. Major assemblies are necessary to substantially

advance the state-of-the-art.


10


STRUCTURAL FLIGHT EXPERIMENTS

• MUST BE FLIGHT WEIGHT STRUCTURE

• MUST BE MAJOR STRUCTURAL ASSEMBLY NOT

SMALL PANELS

• MUST SIMULATE SIGNIFICANT DESIGN AND FUNCTIONAL ASPECTS OF OPERATIONAL AIRCRAFT

GP74-1039-11

MCDONNE ...... AIRCRAFT COMPANY

11


ACTIVELY COOLED PANEL INSTALLED ON ASSET

To support our conviction that small panels are only adequate

to conduct exploratory research, let us review the ASSET program.

An insulated actively cooled panel was flown on ASSET (ASV-3) in

July of 1964. !1aximum speed and altitude were 18,000 ft/sec and

212,000 ft, respectively. The test panel covered an area of 24

inches by 12 inches, was furnished by the Bell Aerosystems Company

and installed at the back under surface of the vehicle as shown

for the flight demonstration by MCAIR.


12


ACTIVELY COOLED PANEL INSTALLED ON ASSET

15000 F

13

BELL ACTIVELY . COOLED PANEL

GP74-1039-91

MCDONNELL AIRCRAFT CONIPANY


BELL ACTIVELY COOLED PANEL EXPERIMENT

The chart shows some of the details of the ASSET actively cooled panel flight

experiment. A l2-inch by 24-inch aluminum panel containing integrally tubed coolant

passages was installed at the aft end of the vehicle as part of the bottom surface

structure replacing the titanium structural floor. A propylene glycol/water mixture

was pumped through the panel during the flight to cool the panel to acceptable

levels. The panels were further protected by an external thermal protection system

as shown. This thermal protection system consisted of two columbium heat shield

panels and a radiation barrier insulation system. The heat shield panels were fab

ricated from a slat array of D-14 columbium alloy members and cross-channels attached

with D-14 rivets and protected with an LB-2 slurry coating.

We suggest that while this experiment essentially demonstrated some basic prin

ciples of an actively cooled structure, it did not convince the critical structural

design engineer of the readiness of active cooling for system application. Perhaps

this is because of: (1) The small scale of the test panel, 2 ft 2 ; (2) the low level

heating experienced by the cooled panel due to the heat shielding; (3) the non

critically of the cooled panel. The vehicle could have survived without the cooling

system functioning; (4) the simplicity of the system compared to an operational

aircraft system; or (5) only one flight and, thus, incomplete investigation of the

total operational environment.

Whatever the reason or combination of reasons it is clear that confidence to

proceed forward with structural technology has historically been gained in a system

atic series of steps with ultimate commitment preceded by a significant structural

demonstration.


14


BELL ACTIVELY COOLED

PANEL EXPERIMENT

RELIEF VALVE

DOUBLE WALL TUBE SHEET - BONDED AND

RIVETED TO STRUCTURAL FLOOR (ALLOW TEMP. 160°F)

RESERVOIR

GEAR PUMP

BYPASS ~ ..

VALVE

a....--------------1 INSULATION PACKAGE .I!;~~~~~~~~~~~0.75

0.003 INCONEL FOI L ,t ';:;= : T COVERING SUBMICRON j "'-0.25 REFRASIL

POWDER (ADL-17) AND FIBERFRAX OUTER WALL INSULATION

SURFACE PANEL (ONE OF TWO SECTIONS SHOWN)

COMPONENT AND COOLING SYSTEM SCHEMA TIC

GP74-103989

MCDONNELL AIRCRAFT COItIIPANY

15


GENERAL COSTING GROUND RULES

The cost analysis was based on the ground rules indicated on

the chart. A minimum management and development experimental shop

philosophy was used to minimize costs. Details of this approach

will be discussed later. It was assumed that the experiments for

the X-24C would be developed such that provisions for the experi

ment could be incorporated in the basic aircraft at initial fabri

cation and thus avoid the costly modification approach.


16


GENERAL COSTING GROUND RULES

• ONE FLIGHT ARTICLE

• LOW COST PROGRAM MANAGEMENT APPROACH

• PROVISIONS IN BASIC AIRCRAFT TO ACCOMMODATE

EXPERIMENT, CHARGED TO EXPERIMENT

• BUDGETARY 1974 DOLLARS GP74-1039-13


17


COST AND SCHEDULE SUMMARY

(Integral Fuel Tank)

Our study indicates that the integral fuel tank experiment

will require 17.5 months from go-ahead to delivery. We have

allowed an additional 7.5 months for installation and checkout

of the experiment at the test site, conducting the flight

test program and evaluating the flight test data. Thus. the

total program span is 25 months. The costs and span time for

each program phase is given in the chart. More detailed descrip

tion of the work elements and costs will be presented later. The

total program budgeting cost is just under $4.9 million given

constant in 1974 dollars. You will note that the management costs,

which represent approximately 5% of the total costs, are listed

as a separate item since they apply to all phases of the program.


18


COST AND SCHEDULE

SUMMARY

YEARS FROM GO-AHEAD ACTIVITY

1 2 3

PROGRAM MANAGEMENT, INTERFACE COORDINATION I AND CONTROL, DOCUMENTATION

DESIGN, DEVELOPMENT AND TEST I

(PHASE I)

TOOLING, FABRICATION AND I

ASSEMBL Y (PHASE IT)

FLIGHT TESTING (PHASE ill) I I

TOTAL COST*

*Constant 1974 Dollars

COST*

274,000

2,055,000

2,318,000

205,000

4,852,000 G P 74-1 039-99


19


COST AND SCHEDULE SUHMARY

(Active Cooling System)

Twenty-four months are required from go-ahead to

delivery of the active cooling system experiment. The

design and development of the heat exchanger is a sig

nificant time pacing item. We have allowed 7 months for

the test phase which 'results in a total program span of

31 months. The total,program budgeting cost is approxi

mately $5.8 million given in constant 1974 dollars.

MCDONNEL.L. AIRCRAFT COMPANY

20


COST AND SCHEDULE

SUMMARY

ACTIVITY YEARS FROM GO-AHEAD

COST* 1

PROGRAM MANAGEMENT, INTERFACE COORDINATION AND CONTROL, DOCUMENTATION

DESIGN, DEVELOPMENT AND TEST (PHASE I)

TOOLING, FABRICATION AND ASSEMBL Y (PHASE II)

FL I G HT TESTI NG (PHASE :m)


21

2 3

I 326,000

1 3,397,000

1 1,856,000

1 198,000

TOTAL COST* 5,777,000

GP74-1039-9

_CDONNELL AIRCRAFT CO_PANY

"tJ :D o Cl :Dr }>o $:~ mC"l rO mC/l $:-1 m Z -I C/l


LOW COST PROGRAM ELEMENTS

Achieving the best product at the lowest cost is a utopia seldom achieved.

Minimizing costs generally results in either reducing the specified capability

required of the product or accepting the risk that the product will not perform as

desired. If this were not axiomatic then the new low cost approach would be the

new norm from which one would again want to reduce cost. Thus the low cost program

elements we have listed in the charts must be considered relative elements, i.e.,

an approach which will provide risks acceptable for an experimental research program

but not acceptable for a high production program. Specifics of each item will be

discussed in the following charts.

ItIICDONNE ...... AIRCRAFT COItIIP'ANY

22


LOW COST PROGRAM ELEMENTS

• RESPONSIBLE MANAGEMENT

• NO EXTRANEOUS CONTROLS

• DEFINITIVE DRAWING SYSTEM

• AUSTERE GROUND TESTING

• EXPERIMENTAL SHOP

• MINIMAL CONTRACTOR FLIGHT TEST SUPPORT GP74-1039-77

MCDONNEL.L. AIRCRAFT C~PAN'"

23


PROGRAM MANAGEMENT

More reliance must be placed on the ability of a singular individual to make

the right decisions ~vithout the benefit of input from his peers. The program manager

must have the full corporate authority to make decisions on the spot which affect all

elements of the program, technical problems, schedule and cost. The personnel assigned

to such a program should be the more experienced people who work at a higher rate of

pay than the average individual but require little supervision thus increasing the

overall cost productivity.

Adhering to a crisp schedule is a must--time is money.

A centralized project location wherein the engineering procurement and manufacturing

team are located in close proximity cuts down the lines of communication and increases

direct communication to resolve problems in an expeditious manner.


24


PROGRAM MANAGEMENT

• CORPORATE AUTHORITY FOR DECISIONS

• CRISP PROGRAM SCHEDULE

• CENTRALIZED PROJECT LOCATION

GP74-1039-78


25

---- ------


PROGRAM CONTROL

Each of the controls that are negated on the chart are part of the normal

systematic means of assuring a quality end product. The larger more complex the

product, the more important such controls are. How then do we propose to assure

reasonable quality and dependability if all controls are removed.

The answer is to place a great burden on the integrity and expected performance

of the contractor. Each individual, be he engineer or assembly man, must accept the

responsibility of doing his job right the first time. Close coordination with all

vendors is essential. The customer/contractor coordination must take on a routine

rather than periodic nature to assure current visibility to progress and problems

and rapid decision making.


26


PROGRAM CONTROLS

• NO MANAGEMENT INFO SYSTEM

• NO PIECE PARTS CONTROL

• NO ZERO DEFECTS PROGRAM

• NO MAINTAINABILITY/RELIABILITY PROGRAM

• NO QUALITY ASSURANCE CAD'S

• NO MIL SPECS - DRAWINGS SHOW REQUIREMENTS

• CLOSE CUSTOMER/CONTRACTOR COORDINATION GP74-1039-79

MCDONNEl.l. AIRCRAFT COMPANY

27


DRAWINGS

The experiment program would be a one of a kind endeavor, with close teaming of

all the functional groups permitting a greatly simplified drawing system. Detail parts

and subassemblies could all be made from a common drawing rather than requiring a number

of drawings. The skill of the manufacturing people could be used in making the best

installation and in making things fit. Thus the engineer would not have to work out

all the details on the drawing, but would only indicate general requirements and

arrangements on the drawing. Developments of such things as line routing would be made

on the actual hard,..rare from similar minimum development information drawings.

Where needed engineering would provide sketches to clarify details. Minimum

information would be put on the drawing, for example, a field callout could be des

criptive enough to negate the need for a bill of material. Fewer drawings would result.

Checking would be done by the designer not another individual charged with

checking. The burden is greater on the designer to assure that it is done right, the

first time.

The question of using the Cathode Ray Tube in the design/manufacturing process is

rather uncertain. Our experience at MCAIR indicates that even for a one of a kind item

the CRT is a fantastic tool which will reduce costs. Since it would probably require

much data to convince people of this we have chosen to cost on the basis of not using

the CRT. However we believe it is an attractive option.


28

REPORT MDC A3265 20 JANUARY 1975

DRAWINGS

• MANUFACTURING DEVELOPMENT DRAWINGS

• SOME SKETCH TYPE

• MINIMUM INFORMATION - FEWER DRAWINGS

• NO CHECKING

• CRT DESIGN TOOL AVAILABLE

GP74-1039-80


29


DEVELOPMENT AND FLIGHTWORTIIINESS APPROACH

For costing purposes we assumed that all the wind tunnel tests will have been

completed on the basic X-24C program and that no new tests would be required.

The structural verification program would consist of conducting failure tests on

the most critical components. The flight article would then be used (rather than a

separate assembly) and tested to 115% of the design limit load (not the normal ultimate

of 150%) to provide confidence in the load paths and interactions. This combined

approach entails greater risk but should be satisfactory for an experimental aircraft.

Fatigue tests of course are a luxury for this type program. We have assumed that no

mockups, spatial, engineering or functional, would be developed. This approach, at

least on paper, will save money. The functional operation of all subsystem components

integrated in the vehicle will be verified through bench tests as well as in the

assembled sys tem.


30

-I


DEVELOPMENT AND

FLIGHTWORTHINESS APPROACH

• NO WIND TUNNEL TESTS

• CRITICAL COMPONENT STRUCTURAL FAILURE TESTS

• STRUCTURAL PROOF (115%) FLIGHTWORTHINESS TESTS

• NO FATIGUE TESTS

• NO MOCKUPS

• SUBSYSTEM FUNCTIONAL AND OPERATIONAL VERIFICATION

GP74-1039-81

MCDONNELL AIRCRAFT COItIfPANV

31


EXPERIHENTAL SHOP

The experiment will be assembled in an experimental shop with limited access. This shop will be equipped with the tooling required to assemble the airframe, test equipment for checkout of installed systems and proof loading, work benches, storage area, inspection supplies, and basic simple machinery to make simple parts. Sheet metal and machined detail parts will be made in shops within walking distance of the experimental shop. Additionally, we will house the project service personnel there, such as engineers, tool designers, production planners, industrial engineers, production control and direct line management personnel. All assembly functions will be performed by highly skilled experimental mechanics.

All shop orders and instructions will be originated by production planning. Normal channels of paperwork will be cut to basics only, with orders and written instructions hand carried directly to those performing the actual operations.

Assembly tooling will consist of the fewest possible tools containing only those features required to control contours and physical interfaces, maintain critical dimensions and locations of structural components to the accuracy specified in the design of the aircraft. The durability will be sufficient to properly build one aircraft.

Machined details will be made on standard machines as far as economically practicable. In all cases, special holding and positioning fixtures will be avoided and depend on setups and simple clamping devices augmented with tracer guided cutting for the machined parts.

Tool design will use sketch type drawings containing only essential information backed up with a liaison function.

All shop orders will be printed on a distinctive color paper to act as an easily recognized flag. This flag will signify a priority for all operations and handling.


32


EXPERIMENTAL SHOP

Q WOOD, FIBERBOARD TOOLING WITH "C" CLAMPS AND TEMPORARY

HOLDING DEVICES

CI 20 TO 25% SUBASSEMBLY TOOLS

• SHOP LAYOUTS IN LIEU OF TEMPLATES

• SHOP ORDERS FOR TOOLS, DETAIL PARTS AND ASSEMBLIES

• SIMPLIFIED PARTS ROUTING

• NO SPECIAL CONVENIENCE TOOLS - TRIM AND HAND CUT

tI ENGINEERING/MANUFACTURING/INSPECTION TEAMING

GP74-' 039-82


33


FLIGHT TEST SUPPORT

It was assumed that the major burden for the flight test program would be borne

by the customer.

The experiment contractor would provide technical support for the installation and

checkout of the experiment in the X-24C aircraft and supply knowledgable personnel for

assistance and consultation during the flight operations.

The customer would be responsible for all the test planning, conducting the

tests, analyzing data, evaluating results and preparing whatever reports deemed

necessary.


34


FLIGHT TEST SUPPORT

• EXPERIMENT CONTRACTOR PROVIDES TECHNICAL SUPPORT FOR INSTALLATION AND CHECKOUT

• GOVERNMENT CONDUCTS TESTS, REDUCES AND ANALYZES DATA, WRITES FINAL REPORT

• CONTRACTOR AVAILABLE FOR CONSULTATION

GP74-1039-83


35

Z -t m

-t Cl l>:O zl> ",r

m!: XP "C m:00 ~I m-< ZO -t:O o

Cl m z


INTEGRAL LIQUID HYDROGEN TANK EXPERIMENT

A number of studies have been conducted to evaluate liquid hydrogen tank structural

arrangement for hypersonic cruise aircraft. Fuselage/tank structural arrangements that

have been proposed for these vehicles consist of either integral or non-integral tanks

with circular or intersecting-circle type (multi-bubble) cross sections. Circular cross

section cylindrical tanks, both integral and non-integral, appear to offer reasonable

vehicle performance and are nearly state-of-the-art designs. The analytical studies

indicate, however, that structural weight and volume utilization advantages may be

attained using the relatively undeveloped integral multi-bubble concept for vehicles

with non-circular fuselage cross sections.

The objective of this experimental program would be to design and develop such a

tank and flight test it aboard a generic version of the X-24C aircraft. The fuselage/

tank structural system would be installed, with the required functional LH2 system

components, by interchanging it with the basic aircraft payload bay. The basis for this

conceptual design study is the -12F version of the X-24C, geometric and flight

characteristics of which were provided by NASA.


36

/

/


HyDRO ........

\ \

\ \

\ \ \ "', \ )

I I

( TANK

'~\ ------------ )---------- GP74-1039-26

MCDONNE ......

37

A'RCRAFTC OIIIfPANY


EXPERIMENT PROGRAM OBJECTIVES

The specific objectives of this program are to develop a flight weight, flight

worthy, thermostructural system and associated subsystems and then to demonstrate

the integrated system within the total flight environment. This demonstration is

significant since it is nearly impossible to simulate all aspects of such an environ

ment simultaneously on the ground.

The thermostructural system will consist of an integral, multi-bubble, liquid

hydrogen fuel tank, its transition and support structure, and the thermal protection

system. The thermal protection system would not be developed as a part of this .

program, since it is intended to use the identical heat shield system as used on the

basic X-24C and reuse as much as possible of the existing heat shields for the tank

experiment.

The liquid hydrogen functional subsystem will provide the capability to vent,

pressurize, and purge the fuel tank, and to dump fuel.

Demonstration of the integrated system will be conducted throughout the full

flight envelope of the X-24C test aircraft, both for normal and simulated emergency

conditions. Ground operations will familiarize crews with all phases of handling

the system including inspection, maintenance, and cryogenic fuel handling.

MCDONNELL A'RCRAIFT COIMPANY

38



• DEVELOP A FLIGHTWORTHY THERMO-STRUCTURAL SYSTEM

• STRUCTURAL TANK FOR LH2 CONTAINMENT

• TRANSITION AND SUPPORT STRUCTURE

• DEVELOP FLIGHTWORTHY SUBSYSTEMS

• VENT

• PRESSURIZATION

• DUMP

• PURGE

• DEMONSTRATE THE INTEGRATED SYSTEM WITHIN THE TOTAL ENVIRONMENT

• FLIGHT OPERATION

e GROUND OPERATIONS (INSPECTION AND MAINTENANCE) GP74-1039-20 MCDONNELL AIRCRAFT COMPANY

39


DESIGN CONCEPT

The design concept defined for this experiment encompasses the operational aspects of the same systems which would be used in future hypersonic cruise vehicles. Illustrated here is the assembled experiment package as it would appear in the payload bay of the X-24C aircraft.

It is from the tank portion of the thermo-structural system that this experiment derives its name. The tank itself is of a substantial size, being approximately eight feet long. Its multi-bubble design features integrally-machined stiffeners and straightline load carrying elements which are the same type design details currently being proposed for future hypersonic cruise vehicles. Domed tank ends are included for increased volumetric efficiency at minimum weight. Thermally induced changes in tank dimensions relative to the remainder of the aircraft structure are accommodated by motion of the support structure connecting them. Detail fabrication and assembly procedures are similar to those that would be used for the tanks of larger vehicles.

The thermal protection system utilizes the basic X-24C aircraft system. It has been assumed that a major portion of the payload hay heat shields will be reused for the experiment. New shields would he fahricated only for those areas in which access or thermal motion accommodations differ from the basic aircraft.

Access to the tank interior is gained by removing two heat shields, unbolting two of the tank support links, and removing the manhole cover in the aft dome of the center tank bubble. These provisions allow maintenance and inspection of the tank, even with the experiment package installed aboard the X-24C aircraft. Structural integrity of the package will not be impaired by removal of the links with the flight test aircraft static on the ground.

Operational elements of the liquid hydrogen system such as the supply port, vent and dump lines are identified here. Provisions for these lines must be made in the basic X-24C aircraft. The helium bottle located aft of the tank is the source of pressurant and purge gas necessary to function of the system.

NrCDONNELL AIRCRAFT CONrPANY

40

STA 206

--


DESIGN CONCEPT

--- --

INTEGRAL LH2

TANK

TANK ACCESS

LH2 DUMP LINE

41

STA 326

LH2 VENT LINE

GP74-1039-64



OVERALL DESIGN CONSIDERATIONS

As a basis for developing a conceptual design for this experiment certain ground

rules ,.,ere established and assumptions made. A basic ground rule ~.,as that all load

redistribution between the basic aircraft and the integral tank be accomplished within

the experiment. The effect of this ground rule was to eliminate any weight or design

penalties to the basic X-24C aircraft resulting from accommodation of this experiment.

It was assumed that the all welded tank construction would use 2219 aluminum alloy since

extensive materials property data is available. Thus the possible expense of material

characterization testing could be eliminated. The tank was designed to have planar webs

at the bubble intersections and straight line elements to provide a practical fabrication

approach and to minimize structural discontinuities. Access to the tank interior while

the experiment is on board the X-24C aircraft is provided for inspection and maintenance

during the program.

The tank external insulation is purged with dry nitrogen gas to prevent cryopumping.

A major cost saving assumption was that the basic aircraft heat shields would be reused

wherever possible on the experiment installation.

Design of the liquid hydrogen system utilized a gaseous helium pressurant to main

tain a tank pressure of 20 psig, dump the hydrogen in a maximum time of thirty seconds,

and purge the empty tank four times for safety. An important design goal was to maximize

the amount of liquid hydrogen carried in the experiment, within the design constraints

dis cussed above and within the limited design flexibility afforded by the X-24C payload

compartment geometry.


42



• STRUCTURAL DESIGN • LOAD REDISTRIBUTION COMPLETE WITHIN PAYLOAD BAY

• TANK CONSTRUCTED FROM WELDED 2219 ALUMINUM ALLOY

• TANK 8 FT LONG, WITH PLANAR WEBS AND STRAIGHT LINE ELEMENTS

• ACCESS FOR INSPECTION AND MAINTENANCE

• THERMAL DESIGN • N2 PURGED EXTERNAL TANK INSULATION

• MAXIMUM REUSE OF BASIC X-24C HEAT SHIELDS

• LH2 SYSTEM DESIGN • GASEOUS He PRESSURIZATION TO 20 PSIG

• 30 SEC MAXIMUM DUMP TIME

• PURGE EMPTY TANK 4 TIMES

o MAXIMUM LH2 CONTAINMENT

43

GP74-1039-19



STRUCTURAL CONCEPT

Illustrated here is the structural concept of the experiment package. The tran

sition structure will mate with interchangeable splice patterns at the forward and aft

ends of the X-24C payload bay. This structure redistributes the basic aircraft

internal loads into discrete loads at the ends of the support links. Those loads are

then transflitted to the structural (integral) tank through the links. Monoball rod

ends allml1 the links to follow thermally induced relative motion between the tank and

the basic aircraft structure thereby minimizing thermal stresses while maintaining

basic load carrying capability. Dimensional changes in the tank resulting from

chilling it with liquid hydrogen amount to nearly six tenths of a percent.

The tank structure is an all welded construction of 2219 aluminum alloy. The

body of the tank consists of three intersecting cones the shells of which are stiffened

by machined flanges in an "Isogrid" pattern. This concept was developed by the

Douglas Aircraft Company in conjunction with the National Aeronautics and Space

Administration. A web is placed at each bubble intersection to carry shear and to

allow internal pressure loads to be self-reacting at minimum weight. The domed tank

ends are also of machined "Isogrid" construction.


44


STRUCTURAL CONCEPT

/

TRANSITION STRUCTURE

• SPLICES • REDISTRIBUTES LOADS

'- SUPPORT STRUCTURE • TRANSMITS AI RCRAFT LOADS

• ACCOMMODATES THERMAL MOTION (DIMENSIONALLY 0.6%)

TANK STRUCTURE • CARRIES AIRCRAFT LOAD

• CONTAINS LH2 • ISOGRID CONSTRUCTION

GP74-1039-70

• WELDED 2219 ALUMINUM ALLOY MCDONNE ...... AIRCRAFT COMPANY

45


THERMAL PROTECTION SYSTEM

The thermal protection system for the experiment serves the two-fold purpose of

protecting the basic structure (the tank) from aerodynamic heating and minimizing

boiloff of the liquid hydrogen fuel being carried. The heat shields will be basically

those used on the X-24C aircraft. It has been assumed as an economy measure that

seventy-five percent of the payload bay panels can be reused on the experiment with

new shields being introduced where access or thermal deflection criteria differ from

the basic aircraft. These nonstructural panels have a pattern of slotted attachment

holes so that no thermal stresses are induced in them. Air loads on the panels are

transmitted directly to the tank by the titanium support structure.

Tank insulation is laminated from two different types of insulation. The inner

layer, bonded to the tank outer surface, is closed cell, freon blown, polyurethane

foam. This insulation provides a good thermal barrier at minimum weight but would

deteriorate if exposed directly to radiation temperatures from the heat shields.

Therefore a fibrous insulation of the microquartz type, which can withstand these

temperatures, is bonded to the outside of the polyurethane foam. It is, in turn,

covered with a foil thickness layer of perforated aluminum. The fibrous insulation

layer is then purged to 0.5 psi with dry gaseous nitrogen which prevents crvopumping

at minimized insulation thickness.


46


THERMAL

PROTECTION SYSTEM

TANK INSULATION

• MINIMIZES LH2 BOILOFF • PREVENTS CRYOPUMPING • N2 PURGED (0.5 PSIG) • WITHSTANDS HIGH AND LOW

EXPOSURE TEMPERATURE

INSULATION FIBROUS

POLYURETHANE FOAM

47

. .yo] l" J= 5=J ~\- -14 ~ T .75 IN. IN. 1 IN.

HEAT SHIELDS

• SUPPORTED FROM TANK • REUSES 75% OF X-24 SHIELDS • ACCOMMODATES THERMAL

DEFLECTIONS

GP74-1039-63



LIQUID HYDROGEN SYSTEM

The operational functions of this system have been mentioned previously.

They are the same as those that would be incorporated in a larger hypersonic

cruise aircraft with only fuel pumps missing. The X-24c research aircraft

will spend extended time aboard a carrier aircraft so it is necessary that

pylon connections and carrier aircraft provisions provide safe venting and

dumping of the hydrogen under these conditions. It is also desirable that

provisions be available for "topping-off" the hydrogen tank just prior to

launch.

The helium pressurization system was sized to provide the capability

to maintain a constant tank pressure of 20 psig and then to dump the full

hydrogen load in a maximum of thirty seconds and also purge the empty tank

four times.

Fuel outlets located at the lowest point in each bubble of the tank,

gather the fuel at a central point for delivery to the dump lines and masts.

A valved alternate outlet line is also provided so that liquid hydrogen can

be supplied to alternate experiment combinations if desired.


48

He PRESSURIZATION '20 PSIG

30 SEC DUMP 4 PURGE CYCLES

FUEL OUTLETS TO: X-24 DUMP MAST

CARRIER AIC DUMP MAST ALT EXPERIMENT


LIQUID HYDROGEN SYSTEM

VENT AND SERVICING OPERATE EXTERNALLY

OR TO CARRIER AIC

STA 206 ------·l ----.

.. ----.-_------: I ALT. EXPERIMENT LINE

STA 326

DUMP LINE

[/f l} ~--~~r\: \: __ ,.L __ :.:=--_ .. -=-, no

/I"-"~: 'e.

.- : (~ ( I . , : : \ I' ~--- , -.~:

.___ 1

----- : -------.J

49

OPERATIONAL SYSTEM ELEMENTS

NO SPECIAL SUPPORT EQUIPMENT

GP74·1039·61



EXPERIMENT INSTALLATION

By nature of its design this experiment offers simple installa

tion into the x-24c aircraft. After disassembly of the FS 206 splice

the forward fuselage section is hoisted away from the aircraft. The

payload bay is then removed and disassembled and the experiment pack

age is installed by attaching it to the FS 326 interchangeable splice.

The forward fuselage is reinstalled by connecting it at FS 206 and

the aircraft is again a complete structural entity. All systems

leading to or passing through the payload bay must be connected. In

stallation of the heat shields then completes the assembly.


50

STA 206

\


EXPERIMENT INSTALLATION

- --

STD AI RLOG DOLLY

STA 326 I

1. REMOVE FORWARD FUSELAGE 2. DISASSEMBLE PAYLOAD BAY 3. INSTALL TANK EXPERIMENT 4. REINSTALL FORWARD FUSELAGE 5. CONNECT SYSTEMS, VENT AND DUMP LINES 6. INSTALL HEAT SHIELDS

HOISTING BRIDLE

X-24C PAYLOAD

BAY

51

JACK STAND

GP74103965




This chart presents a summary of the integral liquid hydrogen tank experiment

program cost and schedule. A schedule and the accompanying cost is provided for

each of the four major program segments: (1) Program Management, Interface Coordi

nation and Control, and Documentation; (2) Phase I - Design, Development and Test;

(3) Phase II - Tooling, Fabrication and Assembly; and (4) Phase III - Flight Test.

Each of these program segments will be discussed in more detail in later charts.

Total cost of the program is estimated to be approximately 4.9 million 1974 dollars

and it is anticipated that the program would be completed in 25 months.

It is interesting to note that Phase I and II costs are nearly equal and that

together they form more than 90% of the total program costs. Flight test costs,

however, represent only slightly more than 4% of the total with the program manage

ment function comprising the remainder.

NlCDONNELL AIRCRAFT CONIPANY

52


COST AND SCHEDULE

SUMMARY


1 2 3

PROGRAM MANAGEMENT, INTERFACE COORDINATION I AND CONTROL, DOCUMENTATION


(PHASE I)

TOOLING, FABRICATION AND I

ASSEMBL Y (PHASE IT)

FLIGHT TESTING (PHASE ill) I

TOTAL COST*


53

COST*

274,000

2,055,000

2,318,000

205,000

4,852,000 GP74-1039-99


PHASE I - DESIGN, DEVELOPMENT AND TESTING

In the following charts we will review in more

detail the significant elements of Phase I, which

affect the design, development and test costs.

54



PHASE I

DESIGN, DEVELOPMENT AND TESTING

GP74-1039-23

MCDONNELL AIRCRAFT C~PANV

55


NO COST DESIGN INFORMATION/ITEMS

An important part of the cost basis for this experi

ment is the assumption that certain design information/

items are available to the experiment contractor at "No

Cost". The representative list of items presented here

would be expected to be available in final form at the

time of authorization to proceed.


56


NO COST DESIGN

INFORMATION/ITEMS

• STRUCTURAL DESIGN CRITERIA/DESIGN LOADS/ STIFFNESS REQUIREMENTS

• AIRCRAFT TRAJECTORY/AERODYNAMIC CHARACTERISTICS

• DRAWING /TOOLING ADEQUATE TO DESIGN AND CONSTRUCT THE STRUCTURAL SPLICES AT FS 206, 326

• AERODYNAMIC HEATING CHARACTERISTICS FOR HEAT SHIELD DESIGN

• THERMAL DESIGN OF HEAT SHIELDS

• INTERFACE CONTROL DOCUMENTATION

GP74-1039-74

MCDONNEL.L. A'RCRAFT COItIIPANY

57


INTERFACE CONTROL DOCUMENTATION

Interface control documentation in the form of drawings, specifica

tions and reports, containing information adequate to define any special

design requirements, is a vital program element. Generation and utili

zation of this data ensures that both the aircraft and experiment

contractors incorporate proper provisions for the experiment in the

aircraft. The type of information assumed to be available is listed

on this chart.


58

\



INFORMATION ADEQUATE TO DEFINE DETAILS OF:

o TOTAL ENVIRONMENTS - TEMPERATURE, PRESSURE, HUMIDITY, ACCELERATION, VIBRATION, DEFLECTIONS, ACOUSTICAL LOADS

o EXPERIMENTS PAYLOAD BAY - ROUTING/MOUNTING

• COCKPIT DESIGN/CONSOLES - EXPERIMENT CONTROLS INTERFACE

• AIRCRAFT POWER SYSTEMS

• HYDROGEN VENT/DUMP PROVISIONS

o CARRIER AIRCRAFT INTERFACE WITH X-24C

o ON BOARD INSTRUMENTATION SYSTEMS AND PROVISIONS FOR ADDED INSTRUMENTATION

CD FLIGHT TEST FACILITY INSTRUMENTATION SYSTEM GP74-1039-73


59


DESIGN, DEVELOPMENT AND TESTING TASKS

Six major program tasks are identified here for the design, development and testing phase of the integral liquid hydrogen tank experiment. A brief description of each task was written in order to provide a proper basis for cost estimation.

Task 1 - Program Management, Interface Coordination and Control, and Program Document~tion - extends throughout the program. Personnel assigned to this task would be the prime contact to provide NASA with visibility to the program. In addition to technical management and control of budgets and schedules, their responsibilities would include coordination and control, both internal and external, of interface requirements and program documentation.

Task 2 - Preliminary Design - starts with the "no-cost" information provided at authorization to proceed and culminates, three months later, in the design freeze of a final design concept for the experiment. This concept would be arrived at by means of a series of layout drawings and trade studies, backed by technical analysis, aimed at optimization of the experiment to meet program objectives.

Task 3 - Procurement Specifications - would be written and released for those elements of the experiment which were determined, at the time of design freeze, to be more economically bought than made in-house. Technical requirements for flightworthiness qualification of these elements would be included in the procurement specifications.

Task 4 - Shop Drawings - would be created using as a basis the final layout drawings from the preliminary design task. Full release of these drawings would occur twelve months after authorization to proceed. The low cost approach to this design function previously discussed would be followed.

Task 5 - Technical Design Analysis - would be made of the thermo-structural functional systems to analytically verify the design adequacy of the experiment. major efforts would involve strength and thermodynamic analyses.

and The

Task 6 - Development Testing and Planning - would be conducted throughout the predelivery phases of this program. This task involves developing and coordinating the test plans, conducting the tests and evaluating the results.


60


DESIGN, DEVELOPMENT AND TEST TASKS

1. PROGRAM MANAGEMENT, INTERFACE COORDINATION AND CONTROL, PROGRAM DOCUMENTATION

2 PRELIMINARY DESIGN • STRUCTURAL LAYOUTS

• EQUIPMENT AND SYSTEM LAYOUTS

• TRADE STUDIES

• ANALYSIS

3. PROCUREMENT SPECIFICATIONS

4. SHOP DRAWINGS

5. TECHNICAL DESIGN ANALYSIS • THERMODYNAMICS

• WEIGHTS

• STRENGTH • SYSTEM DESIGN

6. DEVELOPMENT TESTING AND PLANNING • DEFINE, SCHEDULE AND COORDINATE ALL TEST PLANS

• COORDINATE, TEST AND EVALUATE - ELEMENT TESTS - SUBCOMPONENT TESTS - PURCHASE PART TESTS

GP74-1039-29

- FLIGHT WORTHINESS TESTS MCDONNELL AIRCRAFT COIt/IPANV

61


TYPICAL ELEMENT TESTING

Element testing is an extremely important portion

of this program phase. It can be used either as part

of development of a design concept or as pre-fabrication

verification of design details. Typical of the element

testing envisioned for this program is the three concept

comparative test of isogrid stiffener splices shown here.

The results of such tests would be used as part of the

basis for selection of the type of splice to be used. For

costing purposes it was assumed that three element tests

of this nature would be required.


62


TYPICAL ELEMENT TESTING

SPECIMENS TESTED TO FAILURE

ISOGRID STIFFENER SPLICE DEVELOPMENT

TAPERED JOINT

I

MECHANICAL ATTACHMENT

63

~ I

1+ + + + J

I WELDED JOINT

I

I •

\ ]1 l I [ { GP74-1 039-1 03



SUBCOMPONENT FAILURE TESTING

The sub-tasks of structural analysis and element testing are only two of the

essential steps envisioned in the approach to demonstration of structural flight

worthiness. Once critical elements have been analytically identified and their

structural capability verified by element test the process will be repeated for sub

components of the experiment structure. Analysis would be used to identify the most

critical of the subcomponents and they will be tested to failure. The purpose of the

tests would be to verify the capability of the subcomponents to sustain ultimate loading.

For purposes of cost estimation three subcomponents were identified, they are a link

with its end fittings, a link-to-transition structure attach fitting, and a tank wall

segment machined to have the final isogrid design.

It is intended that the combination of analysis and failure testing so far

identified be followed by structural proof testing, for both symmetrical and

unsymmetrical conditions, to 115% of design limit load. This proof testing would be

conducted using the completely assembled experiment package and would complete demon

stration of its structural integrity.


64


SUBCOMPONENT FAILURE TESTING

LINK AND FITTINGS

o 4 FT

V

A TTACH FITTINGS

65

ISOGRID

GP74-1039-104



FLIGHTWORTHINESS TESTING

Final flightworthiness testing of the completed experiment would be conducted

in two segments, in plant test and tests conducted at Edwards Air Force Base.

Testing prior to delivery would combine structural proof tests with prelim

inary operational checks of the liquid hydrogen system. There would be two struc

tural tests loading the structure to 115% of design limit load, for the maximum

symmetrical and unsymmetrical conditions. The pressurized liquid hydrogen system

would be exercised simultaneously with the structural test using liquid nitrogen

to simulate the cryogenic fuel. In this manner the systems operational aspects

under deflected geometry may be verified.

After delivery the liquid hydrogen system would again be tested without the

superposition of structural loading but using liquid hydrogen. A final check

for leakage and operation of the hydrogen system would be conducted after in

stallation of the experiment package in the flight test aircraft.

If it is determined by NASA that a ground vibration test of the experiment

is required it would be 'conducted on board the X-24C aircraft at Edwards Air

Force Base. This test, however, is not included in either the schedule times or

cost of this program.


66


. - ~-.-.. - - -_ ... - - ----_ .. _. -_ .. '

FLIGHT WORTHINESS TESTING

PRIOR TO DELIVERY:

• EXPERIMENT TESTED AS A UNIT

• LN2 IN TANK • 2 LOADING CONDITIONS - SYMMETRIC AND UNSYMMETRIC

• 115% DESIGN LIMIT LOADS

• PRESSURE

• CRYOGENIC TEMPERATURES

• LH2 SYSTEM OPERATION

• DEFLECTED GEOMETRY

AFTER DELIVERY:

• LH2 SYSTEM FINAL CHECK

o GROUND VIBRATION TEST

GP74-1039-69


67


STRUCTURAL FLIGHTWORTHINESS VERIFIED

The steps leading to verification of structural flightworthiness

of the experiment package have now all been identified. Careful

analysis would be generated to cover each structural element, both to

verify its adequacy and to identify its relative criticality. Element

testing would then be used to validate the analysis and develop ele

mental design concepts. Failure testing of the three most critical

subcomponents would add confidence that all of the subassembly details

are adequate to carry their calculated ultimate loads. Proof tests of

the assembled package then verify both the capability to carry 115% of

design limit load and internal load distributions. The combined analyt

ical and testing programs would then generate confidence that the X-24C,

with the tank experiment installed, could fly safely to the full extent

of its design envelope.

NlCDONNEI..I.. AIRCRAFT CONIPANY

68


STRUCTURAL

FLIGHTWORTHINESS

VERIFIED

• ANALYSIS

• ELEMENT TESTS TO FAILURE

• SUBCOMPONENT TESTS TO FAI LURE

• PROOF TESTS TO 115% LIMIT LOAD

{). AIRCRAFT CLEARED TO FULL FLIGHT ENVELOPE

GP74-1039-71


69


LH2 SYSTEM COMPONENT QUALIFICATION

It is anticipated that most of the components of

the liquid hydrogen system would be procured items. As

such they would be qualification tested as individual

components by the subcontracting vendor. That qualifi

cation testing, which includes function under predicted

service conditions and component ultimate strength,

would qualify the components for assembly into the ex

periment package system without further testing by the

experiment contractor.


70


~'\ (~?-j

LH2 SYSTEM COMPONENT

QUALIFICATION

• PROCURED ITEMS: BOTTLES, VALVES, SEALS & LINES

• TESTED BY VENDOR FOR: • FUNCTION UNDER PREDICTED ENVIRONMENT • STRENGTH

• COMPONENTS QUALIFIED ON RECEIPT

GP74-1039-30

IMCDONNELL AIRCRAFT COIMPANY

71


PHASE I - DESIGN, DEVELOPMENT AND TEST COST .;ND SCHEDULE SUMMARY

In this chart the Phase I total cost of nearly 2.2 million dollars

is broken down into individual costs for eacll of the tasks previously

identified. Costs are given both in manhours and dollars. The schedule

indicates the seventeen month time span for this phase and critical

milestones such as design freeze and 100% release of the shop drawings,

which are respectively at three and twelve months after authorization

to proceed.

Comparison of costs shown here shows total engineering costs to be

only slightly greater than the cost of testing with the combination

comprising the bulk of the cost.


72

--~~~----------------------------~~~--~--- ------


PHASE I DESIGN, DEVELOPMENT AND TEST COST AND SCHEDULE SUMMARY

ACTIVITY MONTHS FROM GO-AHEAD 1234 5 6 7 8 910111213141516171819202122232425~2728293031 MANHOURS

ATP .()

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL, PROG RAM DOCUM ENTATION ------ -- --h=:::~=:::~=:::~=:::~=:::~;::j

TASK 2 - PRELIMINARY DESIGN ______ -,. DESIGN FREEZE ________________________ j __ j}

TASK 3 - PROCUREMENT SPECS _______ ---- --1>' 90%

TASK 4 - SHOP DRAWINGS ______________ _ ____ ~ (')....LL..,().100%

TASK 5 - TECH_DESIGN ANALYSIS ______ t~~:::::::*=*~~~~ TASK 6 - DEVELOPMENT TESTING AND PLANNING _________________________________ ~~=*~**=*~~£;>)o

TEST PLANS ------------------------------+--- -'m>' ELEMENT & SUBCOMP_ TESTS_______ _ __ _ __ h:~~O)o

PURCHASE PARTS TESTS ___________ __ _ ____ ____ )

FLiGHTWORTHINESS TESTS __________ J_ J ________________ ~::::n

*Constant 1974 Dollars PHASE I TOTALS

4,930

8,200

900

20,840

11,240

32,250

78,360

COST*

136,000

229,000

25,000

581,000

313,000

907,000

(6,000) CFE

2,191,000 GP7410J99a

I nformation contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b). 5 USC 552. Do not disclose outside recipient organization of U.S. Government.


73


PHASE II - TOOLING, FABRICATION ANT' ASSEl-ffiLY

The low cost ground rules and assumptions used to estimate costs for Phase II

have been discussed previously. It should be emphasized, hmvever, that close

coordination and cooperation between management, manufacturing, quality control and

engineering are absolutely essential during this phase in order that a high quality,

low cost product result. In the following charts we will review in more detail the

significant elements of Phase II which effect the tooling, fabrication and assembly

costs.


74


PHASE n

TOOLING, fABRICATION AND ASSEMBLY

GP74-1039-31


75

REPORT MDe A3265 20JANUARY 1975

TOOLING, FABRICATION AND ASSEMBLY TASKS

The end product of the program is a complete assembl~T of the experiI'lent package

with all subsystems operative and completely checked out. As previously explained,

the liquid hydrogen system will have been filled, pressurized and drained in-plant using

liquid nitrogen as a safety measure. The system will also have been pressurized with

gaseous helium so that "sniffer" leak checks could be made.

This phase of the experiment includes tool design and fabrication, manufacture

of detail parts, assembly into subcomponents and finally into the finished experiment

structural assembly. Installation of the required instrumentation and systems then

completes the experiment package. Manufacturing supervision will be charged with main

taining quality while producing this assembly at minimum cost.

MCDONNELL AIRCRAFT COIMPANY

76


T(Q)OlLING, FABRICATION AND ASSEMBLY TASKS

• TOOL DESIGN

• TOOL FABRICATION

• PARTS FABRICATION

• SUBASSEMBLY

• STRUCTURAL ASSEMBLY

• INSTRUMENTATION INSTALLATION

• SYSTEMS INSTALLATION

• PREPARATION AND DELIVERY

(TASK 7)

77

GP74-1039-28



PHASE II - TOOLING, FABRICATION AND ASSEMBLY COST AND SCHEDULE SUMMARY

Phase II involves a 13-1/2 month effort at a cost of $2.4 million. Included

in this total is that portion of the Program Management cost, attributable to the

Phase II effort. These costs, which are attributed primarily to manufacturing

tasks, have been further subdivided (consistent with the tasks shown in the

previous chart) in order to provide more insight to the cost elements. Material

costs and the procured hardware (CFE) costs are also presented.

An important cost element in this phase is the engineering support of the

manufacturing operation which involves some 21,000 manhours and nearly $600,000.

Close coordination between the manufacturing and engineering functions will be

essential to achievement of the low cost program goals previously discussed. This

will involve clarification of drawings and sketches as problems arise and documen

tation of necessary modifications. Engineering personnel will also be responsible

for the liaison function, that is resolution of inspection "squawks" and timely

recommendations for and approval of repairs.

I nformation contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b), 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

78

IWCDONNELL AIRCRAFT COIWPANY

REPORT MDe A3265 ':10 JANUARY 1975

PHASE n TOOLING, FABRICATION AND

ASSEMBLY COST AND SCHEDULE SUMMARY

ACTIVITY MONTHS FROM GO-AHEAD 123456789101112131415161718192021222324252627281293031 MANHOURS COST*

ATP.(~

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL, PROGRAM DOCUMENTATION ________________ ~::;::;::;::;::;::;*;::;*~

TASK 7 - TOOLING, FABRICATION IA

AND ASSEMBL Y--------------------------- - -- - -~=:=~~::::::=~*~=PI

MATERIALS ORDERED

TOOL DESIGN & FABRICATION ___________ t=;*~~ADI

FABRICATION & SUBASSEMBLY --- ---- - --h=~=:=;~~*l

ALL CFE RECEIVED ___________________ - -- _______ J- _ J- -.(~ FINAL ASSEMBLY, SYSTEMS AND INSTRUMENTATION INSTALLATION ________________________ __ _ ____ _ __ _ ____ _ ____ ->\,.

PREPARATION AND SHII~____________ _1. _ __ _ __ Jo


Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b). 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

PHASE II TOTALS

79

2,460 69,000

79,890 2,318,000

- (198,000)

(44,700) (1,160,000)

(8,418) (197,000)

- (11,000)

(5,612) , (131,000)

- (30,000)

82,350 2,387,000

GP74-1039102



PHASE III - FLIGHT TEST

The third, and final, phase of the integral liquid hydrogen tank program ~l7ould

be flight test of the installed experiment package. After the package is delivered

to Edwards Air Force Base it would be installed in place of the payload bay of the

X-24C flight test aircraft and final checkouts made. Flight testing on a "ride-along"

basis could then commence. Significant cost elements of this phase are reviewed

in the following charts.


80


PHASE m FLIGHT TEST

81

GP74-1039-27



FLIGHT TEST OBJECTIVES

The primary objective of the Phase III flight test program is to demonstrate

the integrated experiment within a total hypersonic flight environment. By doing

so with this representative system, confidence would be gained that similar systems

for larger hypersonic cruise aircraft are practical.

This program would demonstrate that integral liquid hydrogen tanks are practi

cal for cryogenic fuel containment in aircraft and that load redistribution systems

that accommodate thermal motions are a state-of-the-art reality.

By means of appropriate measurements taken during the flight tests, it would

be demonstrated that the thermal protection system chosen for this experiment

functions according to prediction.

Flight testing would also serve to demonstrate all the operational aspects of

a functional liquid hydrogen fuel and pressurization system.

In summary, all of these demonstrations would significantly contribute to the

technology data base and design confidence for this type of structural system.


82



DEMONSTRATE THE INTEGRATED EXPERIMENT

WITHIN A TOTAL HYPERSONIC FLIGHT ENVIRONMENT

C) DEMONSTRATE A FLIGHT WORTHY INTEGRAL LH2 TANK o FUEL CONTAINMENT

• LOAD DISTRIBUTIONS

e DEMONSTRATE A FUNCTIONING THERMAL PROTECTION SYSTEM • TEMPERATURE DISTRIBUTION AND HISTORY

• NO CRYOPUMPING • MINIMIZED FUEL BOILOFF

• DEMONSTRATE AN OPERATIONAL LH2 SYSTEM

• FILL • VENT • DUMP • INSPECT ~ MAINTAIN

NORMAL FLIGHT OPERA TION

EMERGENCY FLIGHT CONDITIONS GP74-1039-67

GROUND OPERA TIONS MCDONNELL AIRCRAFT COMPANY

83


FLIGHT TEST PLAN

The proposed flight test plan would utilize varying parameters such as Mach

number, dynamic pressure, load factor, and hydrogen dump rates to assess the

performance of each of the systems. The thermo-structural system would be sub

jected to the full range of aircraft load factors with fuel quantities varying

from empty to full. The thermal protection system would be subjected to a

gradual buildup of aerodynamic heating rates and the liquid hydrogen system

would have demonstrated its capability to deliver fuel at rates varying from

slow to emergency dump.

The program starts with three familiarization flights to accommodate the

pilot to varying rates of fuel usage. Following flights utilize gradual buildups

in flight parameters and simulated fuel usage rates to accomplish the flight test

objectives. Emergency flight conditions necessitating fuel dump are simulated in

two of the flights.

ItIICDONNELL AIRCRAFT COItllPANY

84

TEST CONDITIONS

FLT DYN LOAD MACH

PRESS. FACTOR

1 4 LOW LOW

2 4 LOW LOW

3 4 LOW LOW

4 5 LOW LOW

5 5 MED 0 6 5 HIGH 0 7 6 LOW 0 8 6 ® 0 9 6 ® G) 10 6 ® 0 11 6 HIGH MAX

12 6 HIGH MAX


FLIGHT TEST PLAN

OBJECTIVES

FAMILIARIZATION, LH2 STAYS ON BOARD

FAMILIARIZATION, SLOW LH2 DUMP, SIMULATE USAGE

FAMILIARIZATION, FAST LH2 DUMP, SIMULATE EMERGENCY

START BUILDUP OF M, G, q, VARY LH2 DUMP RATE

BUILDUP OF G, q, VARY LH2 DUMP RATE


BUILDUP OF M, G, VARY LH2 DUMP RATE




MAX FLIGHT CONDITION - SLOW DUMP RATE

MAX FLIGHT CONDITION - SIMULATE EMERGENCY DUMP

o Gradual Load Factor Buildup ® Gradual Dynamic Pressure Buildup Scheduled Maintenance/Ground Operation Provides Servicing/Maintainability Data

GP74-1039-84

MCDONNEL.L. AIRCRAFT COItIIPANY

85


FLIGHT TEST INSTRUMENTATION

Twenty measurands have been identified as the minimum required to monitor

demonstration of the experiment package performance during flight test. Strain

gages, an accelerometer, thermocouples and pressure transducers would be installed

during final assembly of the experiment. This instrumentation would be used to

monitor the same quantities during ground testing and during flight test. The

instrumentation would be designed to be compatible with the airborne instrumen

tation recorder and telemetry system in the X-24C.

IMCDONNEI..I.. AIRCRAFT COIMPANY

86


Fl~GHT TEST

INSTRUMENTATION

CD STRAIN GAGES -10 - VERIFY OVERALL AIRCRAFT LOADS AND INTERNAL DISTRIBUTION

• ACCELEROMETER - 1 (TRI-AX) VERIFY LOAD FACTORS AND VIBRATION ENVIRONMENT

• THERMOCOUPLES - 6 - VERIFY TEMPERATURES, DISTRIBUTIONS AND HISTORIES

o PRESSURE TRANSDUCERS - 3 - VERIFY TANK AND LINE PRESSURES, MONITOR HYDROGEN FLOW RATES

GP74-1039-76 MCDONNEL.L. AIRCRAFT COMPANV

87


FLIGHT TEST TASKS

Flight test tasks would be performed in conjunction with NASA personnel with

the experiment contractor providing support and analysis only on an "as required"

basis.

Final checkout of the liquid hydrogen system would include fill, pressurization,

leak and dump tests of the system using liquid hydrogen and facilities available at

the flight test site. The experiment would then be installed aboard the X-24C

aircraft and a final leak check would be conducted.

If necessitated by NASA requirements, the flight test aircraft would then be

placed on jacks and a ground vibration test to verify satisfaction of stiffness

requirements would be conducted.

The aircraft would then be ready for further ground crew familiarization and

flight test.


88


FLIGHT TEST TASKS

• LH2 SYSTEM FINAL CHECKOUT

• INSTALL EXPERIMENT IN X-24C

• PERFORM GVT IF REQUIRED

• EXPERIMENT CHECKOUT

• PROVIDE SUPPORT AND ANALYSIS AS REQUI RED

GP74-1039-44


89


BASIS FOR FLIGHT TEST COST

The ground rules and assumptions used for estimating flight test

costs are presented here. They emphasize the minimal role of the

experiment contractor in this phase of the program. NASA would use

personnel and facilities already dedicated to the basic X-24C flight

test program to reduce their cost associated with this experiment.

MCDONNELL AIRCRAFT COIHPANY

90


BASIS FOR FLIGHT TEST COSTS

., FLIGHT TESTS AT EDWARDS AFB

CD X-24C AIRCRAFT FLOWN AND MAINTAINED BY NASA

e TEST FACILITIES PROVIDED BY NASA

o DATA REDUCTION BY NASA

o MINIMUM CONTRACTOR SUPPORT

6) 12 FLIGHTS AT RATE OF 3 PER MONTH

(J DATA COLLECTION BY ON BOARD X-24C SYSTEMS

o LH2 AND ASSOCIATED EQUIPMENT AVAILABLE AT EAFB

o FINAL LH2 SYSTEM CHECK CONDUCTED AT EAFB

o GVT CONDUCTED ON BOARD X-24C IF REQUIRED GP74-1039-32

MCDONNELL AIRCRAFT COIKPANY

91


PHASE III - FLIGHT TEST COST AND SCHEDULE SUMMARY

Minimized contractor participation in this program phase is reflected by its

low cost. A six week installation and checkout period is followed by flight test.

It is anticipated that flight testing could be conducted on a "ride-along" basis

at the rate of three flights per month. The remaining two month period would be

utilized in analysis and evaluation of flight testing results. Phase III would

be completed in 7-1/2 months.

The chart indicates a subdivision of cost for the various activities in

Phase III with an allocation of approximately 700 manhours per month for contractor

support during the last six months of the program.

IHCDONNELL AIRCRAFT COIHPANY

92


I~"\ PHASE ill (~~2j

FLIGHT TEST COST AND SCHEDULE SUMMARY

ACTIVITY MONTHS FROM GO-AHEAD 123456789101112131415161718192021222324252627281293031 MANHOURS COST*

ATP .oC~

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL, PROGRAM DOCUMENTAT�ON _________ - -T ---- ---- ----- --- ----- ----TASK 8 - FLIGHT TESL _________________ -1---- _L ----- --- -- ----- ----- ,

DELIVERY OF EXPERIMENT. _______ ---- ---- - --r- ---- - ---- -11--)- J-kl LH2 SYSTEM FINAL CHECKOUT

AND INSTALLATION IN X-24C _____ ---- - -r ---T -TT -r --r -_cO

:~~;~:~::~~~:::~::~~:~T:::: 1 T: ]TPlP-l---I--r~ >

:::~;:~ ;~:P:~~~~~~I~~::::: 1 : I Il: :rl-_ -_--_-__ -_--_- _-__ -_-.I_-J1_-__ -_ -_~:

*Constant 1974 Dollars Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b), 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

FLIGHT TEST TOTALS

93

2,460 69,000

7,630 205,000

(91,400)

}

700M/H } PER MO. (113,600) AVERAGE

10,090 274,000



PROGRAM COST AND SCHEDULE SUMMARY

In summary, demonstration of the integrated tank experiment within

the full flight envelope of the X-24C would be completed in 25 months at

a cost of 4.85 million dollars. The cost of each program phase, with its

associated manhour expenditure, is indicated here.

IfIICDONNELL AIRCRAFT COIfIIPANY

94


I~'\ ("-~~ [)~

PROGRAM COST AND SCHEDULE SUMMARY (~--ACTIVITY MONTHS FROM GO-AHEAD MAN HOURS COST*

1 2 3 4 5 6 7 8 9 10 1112 13 14 15 16 17 18 1920 21 22 23 24 25 26 2/28 29 30 31

ATP.( ).

TASK 1 - PROGRAM MANAGEM~NT, INTERFACE COORD AND CONTROL, PROGRAM DOCUMENTATION _________ ,. 9,850 274,000

PHASE I - DESIGN, DEVELOPMENT AND TEST __________________________________ ). 73,430 2,055,000

PHASE II - TOOLING FABRICATION AND ASSEMBLY __________________________ lI\ 79,890 2,318,000

__ 1. .1 _ _ J_ PHASE III - FLIGHT TEST .---------- _____ ,. 7,630 205,000 - -- - -- - - -- - - -

*Constant 1974 Dollars PROGRAM TOTALS 170,800 4,852,000

A DEMONSTRATED LH2 SYSTEM/TANK INTEGRATION WITHIN ITS TOTAL Information contained h.erein is privileged HYPERSONIC ENVIRONMENT or conf,dent,al InformatIon of McDonnell Douglas Corporation and exempt from pub· lic disclosure under subsection (b). 5 USC 552. Do not disclose outside recipient organization of U.S. Government. 95



CRITICAL TECHNOLOGIES AND LONG LEAD ITEMS

Assessment of the conceptual design and test program indicates that

no technology areas or long lead time procurement problems exist which

would indicate potential program delays. The assumed availability of

2219 aluminum plate for construction of the tank, however, could be

critical to both cost and schedule since current availability is quite

limited.


96


CRiTICAL TECHNOLOGIES

AND lONG lEAD ITEMS

TECHNOLOGI ES - NONE

LONG LEAD ITEMS - NONE

MATERIALS: 2219 ALUMINUM ASSUMED AVAILABLE FOR TANK

GP74-1030-41


97


X-24c PROGRAM IMPACT OF EXPERIMENT ACCOMMODATION

Several areas of the basic X-24C aircraft program deserve special consideration in that

early attention to these details would minimize the impact of accommodation of the integral

liquid hydrogen tank experiment.

Careful attention must be paid, for instance, to design of the FS 206 and 326 struc

tural splices so that they are readily made interchangeable with simple tooling. In areas

outside the payload bay space, routing and mounting provisions must be made for vent and

dump line routing and the dump control itself. Design of the aircraft controls routing

through the bay must consider eventual installation of the tank experiment. If the tank

is to be used as a liquid hydrogen source for alternate or combined experiments, plumbing

of the hydrogen from the tank to the point of use must be considered. In addition, cost

savings could be incurred by considering access requirements to the tank experiment in

design of the basic X-24c thermal protection system heat shields. Requirements for data

collection during the tank experiment flight tests should also be considered in design of

the basic X-24C instrumentation system.

Special tools will be required for installation of the experiment. They will include

tooling, supplied by the aircraft contractor, for the FS 206 and 326 interchangeable

splices, vent and dump masts, and supports for such items as controls, electrical wiring,

and plumbing. The cost of such tooling has not been included in the experiment program

estimates. These requirements should be considered early in the X-24C program.

Perhaps the most important single requirement resulting from accommodation of the

tank experiment is the necessity for the X-24c-12F weight and C.G. envelope used for

design to include the effects of this experiment. Failure to do so could require con

siderable ballast during the flight test program in order to result in a controllable

aircraft.

IMCDONNELL AIRCRAFT COIMPANY

98


X-24C PROGRAM IMPACT OF

EXPERIMENT ACCOMMODATION

• INTERFACES AND STRUCTURAL PROVISIONS • FS 206, 326 SPLICES

• VENT AND DUMP LINE ROUTING

• DUMP CONTROL ROUTING AND COCKPIT INSTALLATION

• CONTROLS ROUTING IN PAYLOAD BAY

• LH2 ROUTING FOR ALTERNATE EXPERIMENTS

• PAYLOAD BAY SHINGLES (COMMONALITY)

• INSTRUMENTATION

• SPECIAL TOOLS • FS 206, 326 SPLICES

• VENT AND DUMP MASTS

• CONTROLS, ELECTRICAL AND PLUMBING SUPPORTS

• BASIC AIRCRAFT DESIGN • WEIGHT AND C.G. ENVELOPE

99

--- --- -----

GP74-1039-43


VERSATILE TEST BED

The integral liquid hydrogen tank experiment has great potential aboard the

X-24C as a test bed for alternate and combined flight experiments. It could

readily be used as a source of cryogenic hy~ogen to be used as fuel for an

experimental scramjet engine or as a heat sink for an active cooling system.

The support structure for the X-24C heat shield is adaptable to a number of

alternate thermal protection systems and the liquid hydrogen tank could be the

basis for testing a variety of external insulative concepts as well as internal

vapor barriers. This adaptability should make possible the "ride-along" flight

test program which has been proposed and reduce overall X-24C flight test program

costs by providing a ready, on-board, source of liquid hydrogen.


100

-,


• LH2 FUELED SCRAMJET

• ACTIVE COOLING SYSTEMS

• ALTERNATE THERMAL PROTECTION SYSTEMS

• LH2 FUEL TANK • EXTERNAL INSULATION

• INTERNAL VAPOR BARRIER

GP74-1039-42


- 101

» -no r:! -< elm IO -io rno Xr -c -rnz :DC) ~CIl rn-< ZCll -i~

:5:


ACTIVE COOLING SYSTEM FLIGHT EXPERIMENT

A number of investigations have identified active cooling of the airframe structure

as a promising thermo-structural design concept for hydrogen fueled high speed aircraft.

Cooled structure which allows the use of conventional aluminum material offers poten

tial advantages in terms of weight, cost, fabricability, and service life. In systems

of this type, coolant is pumped to the cooled aircraft surfaces, to absorb the aero

dynamic heat, and then is returned to a coolant-to-hydrogen heat exchanger. Heat is

rejected to the hydrogen fuel and the coolant recirculated to the surface panels.

Active cooling systems, involving a complete flow distribution system with the

required mechanical components and electronic controls, are susceptible to failures.

System failure could result in a catastrophic structural failure due to exceeding the

temperature capability of the material. Thus, a method or approach to providing "fail

safe" operation is required.

The usual approach to fail-safe operation of cooling systems is through redundancy.

However, full redundancy in a cooling system of the type under consideration is not

necessarily fail-safe and certainly not lightweight. An alternate approach to pro

viding fail-safe operation is to provide means to detect a failure and abort the flight

to achieve safe flight conditions. This is the approach considered in the design of

the active cooling system flight experiment.


102


ACTIVE COOLING SYSTEM

FLIGHT EXPERIMENT

103

GP74-1039-3


-~-- --



Actively cooled structure for hydrogen fueled high speed aircraft shows poten

tial for significant reductions in weight-size-cost of the aircraft. However, the

systems raise concern over risk of cooling system failure, and with it loss of

structural and aircraft integrity. Thus, our experimental objectives are: develop

a flightworthy fail-safe thermo-structural system; develop flightworthy functional

subsystems; and demonstrate the integrated system within the total environment.

The first of these objectives can be met by development of actively cooled

aluminum structure with the inherent capability to survive a cooling system failure.

The second objective can be met by development of an active (convective) cooling

system with circulating coolant and hydrogen as a heat sink. and by development of

a failure detection system to provide the pilot an indication to take corrective action

in case of cooling system failure. The third objective can be met with a flight

test program allowing investigation of normal flight operation, abort flight

conditions, and ground operations.


104



DEVELOP A FLIGHTWORTHY FAIL-SAFE THERMO-STRUCTURAL SYSTEM

• ACTIVELY COOLED ALUMINUM STRUCTURE

DEVELOP FLIGHTWORTHY FUNCTIONAL SUBSYSTEMS

• ACTIVE COOLING SYSTEM WITH HYDROGEN HEAT SINK

• FAILURE DETECTION SYSTEM WITH PILOT INDICATION TO TAKE CORRECTIVE

ACTION

DEMONSTRATE THE INTEGRATED SYSTEM WITHIN THE TOTAL ENVIRONMENT

• NORMAL FLIGHT OPERATION

• ABORT FLIGHT CONDITIONS - "FAIL-SAFE"

o GROUND OPERATIONS - SYSTEM CHECK- OUT AND MAINTAINABILITY

GP74-1039-4


105


LOCATION OPTIONS FOR COOLED STRUCTURE TEST

Five locations on the X-24C research aircraft were considered as viable

candidate locations for the fail safe actively cooled structure experiment.

Considerations were: (1) the test location must provide a significant surface

area; (2) the test location must provide reasonable simulation of the typical

range of heating rates experienced by Mach 6 aircraft concepts, (3) the exper

iment must integrate with the basic aircraft with a minimum cost impact; and

(4) the location should not interfere with the integration of other potential

experiments. On the basis of these factors, the centerline vertical stabilizer

was selected as the best overall location for the test of actively cooled structure.

ItIICDONNEI...I... AIRCRAFT COItllPANY

106


LOCATION OPTIONS FOR COOLED

STRUCTURE TEST

• NOSEWHEEL WELL DOORS

• PAYLOAD BAY LOWER SURFACE

• WING STRUCTURE

• OUTBOARD VERTICALS

• CENTERLINE VERTICAL STABILIZER

SELECTED ~ GP74-1039-90


107



The centerline vertical stabilizer offers a significant structural compo

nent for demonstration of actively cooled structure. The experiment structural

system is designed to be interchangeable with the basic "hot" stabilizer and

also to adapt and utilize the basic X-24c speed brake. In this way costs will be

minimized. The actively cooled structure is designed to operate at an average

surface temperature of 250oF, representative of actively cooled, hypersonic cruise

aircraft concepts. The fin leading edge selected is a low drag cooled concept.

The active cooling system is configured with an efficient coolant utilizing

hydrogen as the heat sink. The integrated system includes a failure detection

system, which will provide pilot warning of malfunction. The structure must be

capable of surviving a tr~nsient heat pulse during abort, without requiring

refurbishment.

IWCDONNEL.L. AIRCRAFT COIWPANY

108

STRUCTURAL DESIGN



e INTERCHANGEABLE ACTIVELY COOLED FIN

e ACCOMMODATES BASIC AIRCRAFT SPEED BRAKE

THERMAL DESIGN

o ACTIVELY COOLED STRUCTURE AT AVERAGE SURFACE TEMPERATURE OF 250°F

o SHARP ACTIVELY COOLED LEADING EDGE APPLICABLE TO HYPERSONIC CRUISE

AIRCRAFT

• ACTIVE COOLING SYSTEM WITH AQUEOUS METHANOL COOLANT AND HYDROGEN

HEAT SINK

Q FAILURE DETECTION SYSTEM TO PROVIDE WARNING OF FAILURE/MALFUNCTION

TO PILOT

o STRUCTURE CAPABLE OF SURVIVING ABORT HEATING GP74-1039-5


109


DESIGN FEATURES OF CONCEPT

This chart illustrates some of the design features of the integrated active

cooling concept. The interchangeable actively cooled fin structure must interface ~vith

the basic fin mounting structure and the basic speed brake.

Provisions must be included in the basic design of the aircraft for the required

subsystem interfacing. This will require space for routing of coolant lines between

the fin and payload bay, space for the failure detection system (FDS) control unit,

environmental control of the control unit, space for wiring between the fin and payload

bay and between the payload bay and cockpit. Aircraft power must be supplied to the

FDS and the cooling system. The illustrated equipment and fin installation provides

insight as to the aircraft areas affected by accomodation of the actively cooled struc

ture experiment. Interface coordination and control will be an important facet of the

experiment programs.

NfCDONNELL AIRCRAFT CONfPANY

110


DESIGN FEATURES OF CONCEPT

., MOUNTING AND ROUTING PROVISIONS INCORPORATED IN BASIC AIRCRAFT

• ACTIVE COOLING SYSTEM - SELF CONTAINED UNIT IN PAYLOAD BAY

ACTIVELY COOLED

ACTIVE COOLING SYSTEM

WIRING TO COCKPIT

STRUCTURE l __ --, FAILURE DETECTION /1

SYSTEM CONTROL UNIT .,,-- ~l

.... ! .--_. ~--<-\-- -- \ ;;;§§F-§~?~~~':{-: /\(\ '. r--~------ --.-~~\

:::: \. (' _ /'" i . \ ---- ( y'" --' \

./? .~ ~. " .~-------~:l-'-- ... -~~"' 1 ./ _/ /'" , ••• _--;.. .J'

.r"\ - /"'''' 'v·-· - .. ,--- \ ,-_------"" ... ------- ~/ ~ _/"'''' i\ _.-~~:-""-"'-"'-"'--;-- i_fJ' ~".".--- ' _____ -_-- _=-:-:".=-.d--<.-~~--------------:-:~~' ?~~;::;:::;.;.-;;.--

c..-----.... I ... ..., __ ( __ J ______ ---------------------------------------- -------------------------------------- ----------.. ----------------------------- ( . .' .... H Y DR 0 G E N

PAYLOAD BAY COOLANT VENT/DUMP

COOLANT RETURN SUPPLY LINE LINE

GP74-1039-6

MCDONNEL.L. AIRCRAFT COItIfPANY

111


FIN GENERAL ARRANGEMENT

The general arrangement of the actively cooled centerline vertical stabilizer is

illustrated. As shown, the fin is basically a wedge slab design \"ith a low drag swept

leading edge. The leading edge diameter is 3/8 inch and the sweep angle 50 degrees.

The wedge total angle is 10 degrees. The fin is configured to utilize the hot speed

brake, the speed brake actuator, and hinge-line seals of the basic aircraft vertical

stabilizer design. The fin is designed to attach to the aircraft fuselage structure by

9 lug type attach points and will be interchangeable with the basic fin design.

MCDONNELL AIRCRAFT CONfPANY

112


FIN GENERAL ARRANGEMENT

3/8 IN. DIA (COOLED) LEADI NG EDGE

6.17 FT

2.61 FT

~

100

WEDGE ANGLE

113

BASIC SPEED BRAKE

BASIC ACTUATOR AND HINGE LINE SEALS

INTERFACE WITH BASIC AIRFRAME

GP741039·7

MCDONNELL AIRCRAFT COIMPANY


FIN STRUCTURAL DESIGN CONCEPT

This chart illustrates some of the details of the actively cooled fin structure.

Master tooled, lug type, attach points, as shown, are a representative means of

achieving an interchangeable interface with the basic airframe structure. A root

closure torque box is provided for redistribution of lug loads.

The fin structure consists of aluminum honeycomb sandwich panels, mechanically

attached to aluminum spars and ribs. The panels are fabricated from an alundnum

outer skin, a non-perforated aluminum honeycomb core and an aluminum inner skin. The

coolant tubes and manifolds are brazed to the outer skin. The outer skin/tube

assembly, the honeycomb core, and the inner skin are bonded together with FM 400

adhesive. The coolant manifolds are designed to serve as edge closures for the honey

comb sandwich. The tubes and manifolds can be assembled and completely bench checked

prior to complete panel assembly.

A high degree of damage tolerance is achieved. The non-perforated honeycomb

core will contain the coolant, providing tubing fail-safe capability. The panel can

sustain significant local overheating. The tube-to-skin bonds act as inhibitors to

crack growth.


114


FIN STRUCTURAL DESIGN CONCEPT

LOCKALLOY

SPARS AND RIBS

FWD MANIFOLD (ALUMINUM) ALUMINUM SKIN

COOLANT TUBE (ALUMINUM)

-$-

~FC2SFFF£+3-------+ I FUSELAGE

(TYP) 9 PLACES

115

GP74-103917 MCDONNE ...... AIRCRAFT COMPANY


FIN THERMAL DESIGN CONCEPT

A typical coolant distribution is shown. Aqueous methanol, se

lected as the coolant, enters a supply manifold at the bottom of the

fin. Alternate tubes carry the coolant to either a manifold at 2/3

span or to the tip manifold. The coolant is then directed to the

leading edge manifold and then returned to the active cooling system

where heat is rejected to hydrogen heat sink via a coolant-to-hydrogen

heat exchanger. The cooled skin panels are of aluminum skin-honeycomb

core with "Dee" shaped coolant tubes. The leading edge is cooled by

the flow of coolant through the integrated leading edge coolant return

manifold. In case of cooling system failure, the outer surface skin

must have sufficient mass to absorb the abort heating. Thus, the skin

thickness is a function of abort heat load.


116


FIN THERMAL DESIGN CONCEPT

COOLANT RETURN

COOLANTTUBE~ MANI:OLDS

--F~L2~-=""~Ll 1'1'1 {fyi\ A-A --..- fisJ llc~ll LLI

LOCKALLOY FWD MANIFOLD LEADING EDGE

DISTRIBUTION MANIFOLD

~ COOLANt

SUPPLY

BASIC SPEED BRAKE

DISTRIBUTION MANIFOLD

THICKENED ALUMINUM SKIN

TYPICAL COOLANT TUBES

GP741039-B

B-B COOLANT RETURN -


117


SUBSTANTIAL AREAS OPERATE AT HEATING RATES ABOVE 5

This chart illustrates the heating rate distribution for I

a typical hypersonic transport at Mach 6 at an altitude of I

105,000 ft. A substantial portion of the actively cooled

surface will experience cruise condition heating rates greater 2

than 5 BTU/FT sec. This indicates that the activel~ cooled

structure associated with an active cooling flight experiment

should be capable of operation within a heating rate environ

ment resulting in rates between 5 to 10 BTU/FT2

sec.

IKCDONNELL AIRCRAFT COIKPANV

118


SUBSTANTIAL AREAS OPERATE ~~\ AT HEATING RATES ABOVE 5 ~

TYPICAL M = 6 CRUISE HST (ACTIVELY COOLED)

UPPER SURFACE

1.5

l

1 t 5

LOWER SURFACE q (BTU/FT2 SEC)

GP74-1039-16


119


NORMAL OPERATING HEATING RATES

This chart illustrates typical heating rate histories

for the cooled centerline.vertical stabilizer. Cruise time

at r1ach 6 is assumed to be 60 seconds and descent from cruise

is assumed to be at maximum aircraft lift-to-drag ratio.

This type of flight profile was selected to insure a cooling

system design of adequate cooling capacity. Increasing angle

of attack above the 5°_6° angle at (L/n) max would tend to

decrease the fin average heating rate and decreasing cruise

time at maximum Mach number would tend to reduce the total

flight heat load and, thus, the total amount of heat sink

required. A number of options are available, as indicated

above, in terms of angle of attack and cruise time. As shown

on the chart, heating rates at the selected cruise condition

vary from values a little above 10 BTU/ft2

sec down to values

below 7 BTU/ft2 sec. This would give simulation of the higher

heating rate areas on a Mach 6 cruise aircraft. Additional

options are available and are discussed in the following

chart.


120

q - HEATING

RATE


NORMAL OPERATING HEATING RATES

DESCENT AT LID = 2.5

10 j-CRUISE AT M = 6, ALTITUDE = 89,000 FT

8~------~~~~-+~------~------~------~

6r------+Hr------~~~~~-T-------~------~

BTU/FT2 SEC 4 r------H-I---t-------+--~~~-----~-----~

2r---~~~-------+--------~~~~~----~

o--~----~------~------~--------~----~~ o 100 200 300 400 500 TIME - SEC GP74-1039-85

MCDDNNE ........ AIRCRAFT COMPANY

121


EXPERIHENT SHfULATES HAXIHUM OPERATIONAL HEATING RATES

The heating rate distribution exhibited is typical for the centerline

vertical stabilizer at Mach 6 and an altitude of 89,000 ft. Dynamic pres

sure at this condition is 1000 PSF. Aircraft angle of attack is 5 degrees.

As shown, the heating rates (turbulent flow) vary from 12 BTU/ft2

sec near

the leading edge to about 7 BTU/ft2

sec on the aft lower portion of the

fin. Operation at a dynamic pressure of 500 PSF at lIach 6 (105,000 ft

altitude) would reduce these heating rate values to approximately 50 per

cent of the illustrated values. This will give the capability to investi

gate the range of heating rates from abort 3 BTU/ft2

sec up to the 12

BTU/ft2 sec. Thus, this range of heating rates includes the heating rates

experienced over a major portion of the surface of a Mach 6 aircraft and

experimental simulation would be good within this altitude range. The

actively cooled fin is designed for the heating rates associated with

operation at a dynamic pressure of 1000 PSF.


122


EXPERIMENT SIMULATES MAXIMUM OPERATIONAL

HEATING RATES

M = 6 CRUISE AT 89,000 FT ALTITUDE TWALL = 250°F .

q LEADING.EDGE = 89.3 BTU/FT2SEC (DIA = 3/8 IN.)

CONSTANTq (BTU/FT2 SEC)

HOT SPEED BRAKE

GP74-1039-15

MCDONNELL AIRCRAFT COIfIfPANY

123 \


FLIGHT ABORT CONSIDERATIONS

In order to provide a "fail-safe" capability to the actively

cooled structure a number of factors must be considered in the system

design and operational utilization. A failure detection system,

capable of providing fast response detection of a system failure or

malfunction, must be provided. The feasibility of utilizing an abort

trajectory to reduce descent heating and total descent heat load,

must be verified. An adequate structural capability must be provided

to survive the heating loads during abort with the associated cooling

system failure effects considered in determining the strength capa

bility required of the fin. All these factors were considered in

the conceptual design of this experiment.


124


FLIGHT ABORT CONSIDERATIONS

• FAST RESPONSE FAILURE/MALFUNCTION DETECTION

• ABORT TRAJECTORY TO REDUCE HEAT LOAD

• ADEQUATE HEAT ABSORPTION CAPABILITY TO HANDLE

ABORT HEATING

FAIL-SAFE CAPABILITY GP74-1039-14

MCDONNELL AIRCRAFT' COMPANY

125

._----------


ABORT TRAJECTORY REDUCES DESCENT HEATING

Maneuvering the aircraft during an abort can result

in significant reductions in the abort heat load. This

is illustrated by the comparison of the high lift-to-drag

ratio (2.5) descent, the low LID (0.58) descent at low

angle of attack with speed brakes deployed, and the mini

mum heating descent. The minimum heating descent utilizes

a constant 2.5g pitch up which transitions to a 90° bank

angle - 30 angle of attack condition. For this trajectory,

the descent heat load for a typical location on the center

line vertical fin is only about 11 percent of the high

LID descent heat load. Thus it appears entirely feasible

to significantly reduce total heat input loads by such

traj ectories.


126


ABORT TRAJECTORY REDUCES DESCENT HEATING

q - HEATING RATE

BTU/FT2 SEC

10------~----~----~----------~----~ START ABO~T~TW = 250°F MACH 6 CRUISE

A AL T = 89,000 FT X = 3 FT 8~----~~~4-----~-----+----~----~

~HIGH LID DESCENT /' Q = 951 BTU/FT2

LOW LID DESCENT Q = 219 BTU/FT2

MINIMUM HEATING DESCENT

O~L---~--~~~~~~~~~~~----~

o 100 200

127

300 TIME - SEC

400 500 600 GP74-1039-86

"'CD~NE&'&' AIRCRAFT COMPANY


SPEED AND DESCENT TRAJECTORY DETERMINE SKIN THICKNESS

The capability of a bare unprotected aluminum skin to survive the

heating during the abort descent depends upon the flight speed at abort

initiation and the type of trajectory utilized to arrive at a safe

flight condition. The initial speed influences the magnitude of the

heating rate. The type abort trajectory also influences the abort

heating rates, but of more importance it determines the time required

for descent, and thus, the total abort heat load. If we utilize a

thermo/structural system in which an external aluminum skin is used as

a heat sink the required skin thickness is determined by the total abort

heat load. The chart illustrates these effects for the case where the

initial structural temperature is 250°F prior to the cooling system

failure requiring the abort and a limit on the structural temperature

rise of 200°F. The dominating effect of the type abort trajectory used

is evident. With this type of heat sink (unprotected aluminum skin)

approach, limitations may be required on maximum test Mach number. This

Mach number limitation will be a function of the ability of the aircraft

to pull high g's during abort and the ability to accommodate vertical

fin weight. The fin design cost was based on the use of 0.08 inch thick

skins.

/MCDONNEL.L. AIRCRAFT COMPANY

128


SPE·ED AND DESCENT TRAJECTORY DETERMINE

SKIN THICKNESS

0.8 r-------,.--------,--------r----r------,

HIGH LID DESCENT

0.6 1--------+------4-------J.-..~---~

ALUMINUM FIN DESIGN WITH 0.08 SKIN LOW LID

THICKNESS O.4~--~-~----~--~~DESCENT-~~ I

INCH STRUCTURAL MINIMUM CONSIDERATIONS HEATING

DESCENT 0.2 1------:(--{---+-----~-----+---'\-7oC..--___1

---- - --- - -~ - --- ---- ------

o~-------~--------~~--------~--------~ 2 3 4 5 6

MACH NUMBER GP74·'039-87

_CDONNELL AIRCRAFT CO_PANV

129


COOLING SYSTEM ELEMENTS

The active cooling system was assumed to be contained in the pay

load bay. The system arrangement shown is one of several arrangements

that could supply the required cooling capacity. The illustrated ap

proach utilizes supercritical hydrogen, stored in a vacuum jacketed

tank, as the heat sink. Another approach, using normal boiling point

hydrogen at subcritical pressures and a pump to deliver hydrogen to the

heat exchanger, could also be utilized. It is estimated that either

approach would result in about the same order of magnitude cost. The

major elements of the integrated system exhibited on the chart are the

cooled fin, the coolant distribution lines, the tank containing the hy

drogen heat sink, the coolant pump package, and the coolant-to-hydrogen

heat exchanger. The heat exchanger design concept will fall within the

category of critical technology.


130


COOLING SYSTEM ELEMENTS

ACTIVELY COOLED

FIN

1

I J

~ ____ ~~J~L / - /

\ ~4 ____ ------PAYLOAD-BAY----------~·\

HYDROGEN TANK

\ HEAT SOURCE TO MAINTAIN

PRESSURE I

VALVE AND REGULATOR

PRESSURE-RELIEF VALVE

,...-- ----~ACCUMULATOR

~ ~ \ ~PUMP & MOTOR

I I

HEAT EXCHANGER

131

--

BASICT SPEED BRAKE

I , COOLANT

RETURN

~COOLANT , SUPPLY ,

-----.--.' HYDROGEN

- VENT/DUMP

GP74-1039-2

MCDONNIELL AIRCRAFT COItIIPANY


ACTIVE COOLING SYSTEM SCHEHATIC

This chart shows a representative schematic with all the elements

integrated into a functional system. Mechanical, electro-mechanical

and electronic components are all required. One of the more critical

items, in terms of'required research and development, is the coolant

to-hydrogen heat exchanger. Obtaining the required heat transfer with

out coolant freezing, and designing to accommodate thermal stresses,

while maintaining reasonable weight and volume are prime areas of con

cern. It is estimated that the total integrated lleating rate over the

entire fin surface is approximately 560 BTU/sec, at Uach 6 and a dy

namic pressure of 1000 PSF, and would require a heat exchanger volume

of 3 ft 3 with a weight of llO lb. Approximately 88 lb of hydrogen

heat sink is required. Operation at lower dynamic pressures at Mach

6 or reducing the maximum cruise Mach number would of course reduce

these values.


132

-,

SUPERCRITICAL HYDROGEN AT

15 ATM PRESSURE

PRESSURE RELIEF VALVE

r , , , , L-

VACUUM SHELL TANK


ACTIVE COOLING SYSTEM SCHEMATIC

TO AI RCRAFT POWER -( -=--= = = ::i1 TO PILOT OFF/ON SWITCH ========--,-..,,1

CONTROL

ACCUMULATOR .:::::{:::. II III PUMP/MOTOR ::::I}:: II III

PACKAGE

~-CONTROL

_--t-----~COOLANT

TO FIN

BYPASS VALVE

METHANOL/WATER COOLANT TO HYDROGEN HEAT EXCHANGER

~--·VENT ---PRESSURE REGULATOR AND FLOW CONTROL VALVE GP74-1039-1


133


FAILURE DETECTION SYSTID1

Features of the selected type of failure detection system are

exhibited by this chart. The sensing elements utilize a eutectic

salt mixture. Melting of the eutectic, at the set point temperature,

changes resistance of the element and thereby changes the magnitude

of a voltage signal to the control unit. Dual sensing elements are

used to prevent false warnings due to element shorts. The selected

set point temperature is 265°F. The sensing elements are positioned

at a location where the normal operating temperature is adequately

below set point temperature. Loss of cooling results in temperature

rise, a signal to the control unit, which in turn provides a pilot

warning of loss of fin cooling. The sensing elements run through

the length of the fin and an overtemperature condition at any point

would be sensed. The FDS control unit requires input power to be in

working order and, therefore, would require connection to an essen

tial electrical power supply.


134


FAILURE DETECTION SYSTEM

INCONEL TUBING 0.089 0.0. (0.064 1.0.)

VOIDS BETWEEN TUBING, CERAMIC AND NICKEL CORE ARE SATURATED WITH A EUTECTIC SALT MIXTURE

POROUS ALUMINUM OXIDE CERAMIC 0.0540.0.,0.034 1.0.

COOLANT TUBE

ALUMINUM SKIN -1: ~CONT~ DUAL SENSING

UNIT ELEMENTS PI LOT II (SET POINT 265°F)

l-~ \I WARNING L -=: AIRCRAFT POWER

135

GP74-1039-33

IIIICDONNELL ~AFT COMPANY



This chart presents a summary of the active cooling system flight

experiment cost and schedule. As shown, the program cost is propor

tioned to four major program facets. The first is overall program

management, interface coordination and control, and documentation.

The remaining three are: Phase I - Design, Development and Test;

Phase II - Tooling, Fabrication and Assembly; and Phase III Flight

Testing. Each of these facets of the program will be covered in

greater detail in following charts. Total cost of the program is

estimated at approximately $5.8 million. The duration is 31 months.

MCDONNELL AIRCRAFT COIWPANY

136


COST AND SCHEDULE

SUMMARY


1 2 3

PROGRAM MANAGEMENT, INTERFACE COORDINATION I

AND CONTROL, DOCUMENTATION


(PHASE I)

TOOLING, FABRICATION AND 1

ASSEMBL Y (PHASE II)

FLIGHT TESTING (PHASE ill) I

TOTAL COST*


COST*

326,000

3,397,000

1,856,000

198,000

5,777,000

G P74-1 039-9


137


PHASE I, DESIGN DEVELOPMENT TESTING

In the following charts we will review in more de

tail the significant elements of Phase I which effect

the d~sign, development and test costs.

ItIICDONNELL AIRCRAFT COItIIPANY

138


PHASE I

DESIGN, DEVELOPMENT AND TESTING

GP74-1039-35


139



Part of the cost basis is the assumption that cer

tain design information/items are available to the ex

periment contractor at "no cost". The items listed, at

authorization to proceed with the experiment, would be

expected to be supplied in a timely manner.


140



• STRUCTURAL DESIGN CRITERIA/DESIGN LOADS AND STIFFNESS REQUIREMENTS

• AIRCRAFT TRAJECTORY/AERODYNAMIC CHARACTERISTICS

• DRAWINGS/TOOLING ADEQUATE TO DESIGN/CONSTRUCT STRUCTURAL SPLICE WITH BASELINE FIN MOUNTING

., DRAWINGS/TOOLING ADEQUATE TO INTERFACE WITH BASELINE SPEED BRAKE DESIGN

o AERODYNAMIC HEATING CHARACTERISTICS DUE TO SPEED BRAKE UTILIZATION AT HIGH SPEED

o THERMAL DESIGN OF SPEED BRAKE HINGE STRUCTURE AND BASELINE FIN SUPPORT STRUCTURE

., INTERFACE CONTROL DOCUMENTATION GP74-1039-37

MCDONNELL AIRCRAFT COIfIfPANV

141



Interface control documentation, in the form of draw

ings, specifications, and reports, containing information

adequate to define any special design requirements will be

a vital program element. It is an absolute fundamental

that the aircraft contractor, as well as the experiment

contractor, be aware of, and plan for accommodation of, any

planned experiments if minimum cost is to be achieved. The

type information expected to be available is listed on the

chart.


142


INTERFACE CONTROL DOCUMENTATDON

INFORMATION ADEQUATE TO DEFINE DETAILS OF:

• EXPERIMENTS PAYLOAD BAY - ROUTING/MOUNTING

• COCKPIT DESIGN/CONSOLES - ACS AND FDS INTERFACE

• TOTAL ENVIRONMENTS - TEMPERATURE, PRESSURE, HUMIDITY, ACCELERATION, VIBRATION, DEFLECTIONS, ACOUSTICAL LOADS

• BETWEEN FIN AND PAYLOAD BAY

• PAYLOAD BAY

• BETWEEN PAYLOAD BAY AND COCKPIT

• COCKPIT

• AIRCRAFT POWER SYSTEMS

• AVAILABLE HYDROGEN VENT/DUMP PROVISIONS

• CARRIER AIRCRAFT INTERFACE WITH X-24C

• ONBOARD INSTRUMENTATION SYSTEM AND PROVISIONS FOR ADDITIONAL INSTRUMENTATION

• FLIGHT TEST FACILITY INSTRUMENTATION SYSTEM GP74-1039-38

MCDONNELL AIRCRAFT COIfIfPANY

143


DESIGN, DEVELOPMENT AND TESTING TASKS

Six major program tasks are identified here for the design, development and testing phase of the active cooling system flight experiment. A brief description of each task was written in order to provide a proper basis for cost estimation.

Task 1 - Program }lanagement, Interface Coordination and Control, and Program Documentation - extends throughout the program. Personnel assigned to this task would be the prime contact to provide NASA with visibility to the program. In addition to technical management and control of budgets and schedules, their responsibilities would include coordination and control, both internal and external, of interface requirements and program documentation.

Task 2 - Preliminary Design - starts with the "no-cost" information provided at authorization to proceed and culminates, five months later, in the design freeze of a final design concept for the experiment. This concept would be arrived at by means of a series of layout drawings and trade studies, backed by technical analysis, aimed at optimization of the experiment to meet program objectives.

Task 3 - Procurement Specifications - would cooling system and the failure detection system. worthiness qualification of these elements would fications.

be written and released for the active Technical requirements for flight-

be included in the procurement speci-

Task 4 - Shop Drawings - would be created using as a basis the final layout drawings from the preliminary design task. Full release of these drawings would occur fifteen months after authorization to proceed. The low cost approach to this design function previously discussed would be followed.

Task 5 - Technical Design Analysis - would be made of the thermo-structural functional systems to analytically verify the design adequacy of the experiment. major efforts would involve strength and thermodynamic analyses.

and The

Task 6 - Development Testing and Planning - would be conducted throughout the predelivery phases of this program. This task involves developing and coordinating the test plans, conducting the tests and evaluating the results.


144

REPORT MDe A3265 20JANUARY 1975

DESIGN, DEVELOPMENT AND TEST TASKS

1 PROGRAM MANAGEMENT - INTERFACE COORDINATION AND CONTROL - PROGRAM DOCUMENTATION

2 PRELIMINARY DESIGN STRUCTURAL LAYOUTS - EQUIPMENT/ELECTRICAL/SUBSYSTEM LAYOUTS

TRADE STUDI ES - ANALYSIS

3 PROCUREMENT SPECIFICATIONS ACTIVE COOLING SYSTEM - FAILURE DETECTION SYSTEM

4 SHOP DRAWINGS

5 TECHNICAL DESIGN ANALYSIS THERMODYNAMICS - STRENGTH - WEIGHTS - SYSTEM DESIGN

6 DEVELOPMENT TESTING AND PLANNING ELEMENT TESTS - SUBCOMPONENT TESTS - PURCHASE PART TESTS - FLIGHT WORTHINESS TESTS - FLIGHT TESTS

GP74-1039-39


145


PHASE I, D, D & T COST AND SCHEDULE SUMMARY

As shown, Phase I will cover a time span of 23 1/2 months. Total man-hours and

cost for this phase are estimated as 66,590 hours and $3,560,000, respectively. The

cost item shown as CFE is procurement cost for the active cooling system and failure

detection system and total $1,700,000. Additional procurement cost are accrued in

Phase II.

Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b). 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

146



PHASE I DESIGN, DEVELOPMENT AND TEST COST

AND SCHEDULE SUMMARY ACTIVITY MONTHS FROM GO-AHEAD

12345678 910111213141516171819202122232~251262 2812!l 30 31 MANHOURS COST*

ATP ~~

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL, PROG RAM DOCU M E NT AT I 0 N __________ I=*:!=!:=!=;:;::!=!:~;:::;::;::;:~=:;::;:::;:~:::::=~

TASK 2 - PRELIMINARY DESIGN ______ ~ DESIGN FREEZE _____________________________ ,,{>-

TASK 3 - PROCUREMENT SPECS

ACTIVE COOLING SYSTEM__________ )-FAILURE DETECTION SYSTEM ___________ ~)-

EVALUATE PROPOSALS _____________ _ ___ ~}

SE lECT VENDORS____________________ -\---? 9,O~.J~OO% TASK 4 - SHOP DRAWINGS______________ _ __ _

TASK 5 - TECH. DESIGN ANALYSIS___ J-h=~=:=~*::Q'). APL'S RELEASED______________________ _1. -I-{> PLACE PO'S_~ _____________________ ~ _________________ I_.Q.

TASK 6 - DEVELOPMENT TESTING AND PLANNING

TEST PLANS____________________________ }o.

ELEMENT & SUBCOMP. TESTS _____ - -f --- -.I. l i /I I I ~ PURCHASE PARTS TESTS ------------ - -r -i-r - IA

FLIGHT WORTHINESS TESTS ------- - -+ -T -y-~- __ 1 __ - --1-- - - -- - -- -~~D! *Constant 1974 Dollars

Information contained herein is privileged Or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (bl. 5 USC 552, Do not disclose outside recipient organization of U.S. Government.

PHASE I TOTALS

147

5,850

8,740

7,210

10,790

8,210

25,790

66,590

163,000

243,000

201,000

301,000

229,000

2,423,000

( 1,700,000)

CFE

3,560,000

GP74 103940



PHASE II, TOOLING, FABRICATION AND ASSEMBLY

The following charts will present a more detailed revieVl of the elements of

Phase II Vlhich effect the tooling, fabrication and assembly costs.

MCDONNEILIL AIRCRAFT CONfPANY

148


PHASE n

TOOLING, FABRICATION AND ASSEMBLY

GP74-1039-45


149


TOOLING, FABRICATION AND ASSEMBLY TASKS

The sub-tasks identified for the Phase II effort are listed on this chart. The

second phase of the experiment program includes tool design and fabrication, manufacture

of detail parts, assembly into subcomponents, and finally the actively cooled fin

structural assembly. Necessary instrumentation components will be installed during the

manufacturing process. Subsystems will be received and checked. In addition, sus

taining engineering support will be required during this period for timely change of

drawings and suggested repairs or modifications, if required.


150


TOOLING, FABRICATION AND 'ASSEMBLY TASKS

• TOOL DESIGN

• TOOL FABRICATION

• PARTS FABRICATION

• SUBASSEMBLY

• STRUCTURAL ASSEMBLY

• INSTRUMENTATION INSTALLATION

• SYSTEMS INSTALLATION

• PREPARATION AND SHIP

(TASK 7) GP74-1 039-47

MCDONNE ...... AIRCRAFT C~.-ANY

151


TOOLING, FABRICATION AND ASSEMBLY COST AND SCHEDULE SUMMARY

The time span of this program phase is 18 months with completion 24 months after

program ATP. Estimated man-hours are 39,570 and cost of $1,954,000 are accrued

during this phase. Material costs are only $40,000 while procurement cost total

$870,000.

Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b). 5 USC 552_ Do not disclose outside recipient organization of U.S. Government.

152



PHASEll TOOLING, FABRICATION AND

ASSEMBLY COST AND SCHEDULE SUMMARY

ACTIVITY MONTHS FROM GO-AHEAD 12345678 91Dl112131415161718192021~232425~272829JD31 MANHOURS

ATP -<>

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL, P R OG R AM DOC U M E NT AT ION _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ ____ t=;:::;::::::;:::!::;::;=!=:;::!:=;:;::!::::;~:::;::::::;~

TASK 7 - TOOLING, FABRICATION AND ASSEMBLY ____________________________________ ~~::::!:::!=!:=;::;::;:::!=:!::=;~~:::;::;~

" MATERIALS ORDERED _______________________ - -. u

TOOL DESIGN & FABRICATION ____ _ __ _ ____ __ ).

FABRICATION & SUBASSEMBLY .__ _L~~::::!=:!=!:::;::!::!:~~). ALL CFE RECEIVED __________________ ---- ____ - -- - ---- - -- - ---- - -- - . .0-FINAL ASSEMBLY, SYSTEMS AND INSTRUMENTATION

A

INSTALLATION ------------------------ -r --- --T -r -r- - ---- - -- - - I PREPARATION ANDSHIP _________ .__ _ __ _ ____ _ __ _1. ___ 1. _____ .co-

*Constant 1974 Dollars PHASE H TOTALS

3,510

36,060

(9,050)

(7,134)

(4,756)

39,570

COST*

98,000

1,856,000

(40,000)

(235,000)

(165,000)

(870,000)

(110,000)

(15,000)

1,954,000 GP74 103948

Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (bl, 5 USC 552. Do not disclose outside recipient organization of U.S. Government.


153

L


PHASE III, FLIGHT TEST

The third, and last, phase of the experiment programs is flight test of the

integrated active cooling system. After the complete experiment package is received

the tes t aircraft \vill have the basic vertical fin replaced by the experiment fin.

The onboard experiment systems will then be installed and exercised during familiar

ization of flight test personnel. Following charts presents elements of Phase III

which influence program costs.


154


PHASE m

FLIGHT TEST

155

GP74-1039-4~

MCDONNELL AIRCRAFT COItIIPANV



The overall objective of this experimental program is to demonstrate an integrated

"fail-safe" actively cooled structure and active cooling system within the total

flight and operational environment. The stated objective offers important opportunities

for the advancement of hypersonic cruise aircraft if the fail-safe abort approach can

be fully demonstrated. The specific objectives and flight test plans are structured to

assure adequate investigation of these opportunities.


156



DEMONSTRATE AN INTEGRATED "FAIL-SAFE" ACTIVELY COOLED STRUCTURE - ACTIVE COOLING SYSTEM WITHIN THE TOTAL

ENVIRONMENT

• DEMONSTRATE ACTIVELY COOLED ALUMINUM STRUCTURE

• DEMONSTRATE ACTIVE COOLING SYSTEM WITH HYDROGEN HEAT SINK

• DEMONSTRATE FAILURE DETECTION SYSTEM WITH WARNING TO TAKE

CORRECTIVE ACTION

NORMAL FLIGHT OPERA TION

ABORT FLIGHT CONDITIONS

GROUND OPERA TIONS GP74-1039-50


157


FLIGHT TEST PLAN

This chart illustrates a suggested approach to meeting the flight test objectives.

The maximum Mach capability would be dependent upon the fin capacity to survive abort

heating. Varying the skin thickness of the fin to increase start of abort Mach number

would have very little impact on program cost. It is possible, with better definition

of the aircraft lee-side flow field, that skin thicknesses on the order of 0.080 in.

could survive a Mach 6 abort.


158

FLT

1

2

3

4

5

6

7

8

9


FLIGHT TEST PLAN ~,

-TEST CONDITIONS TEST CONDITIONS

M = 2 - HIGH DYN PRESS. ACS FUNCTIONAL DEMONSTRATION AND DEMONSTRATE STRUCTURAL INTEGRITY

M = 3 - MED DYN PRESS. ACS AND FDS FUNCTIONAL DEMONSTRATION MANEUVER ABORT TECHNIQUE DEMONSTRATION

M = 4 - MED DYN PRESS. ACS FUNCTIONAL DEMONSTRATION




M = 6 - HIGH DYN PRESS. ACS FUNCTIONAL DEMONSTRATION

M = 6 - HIGH DYN PRESS. ACS AND FDS FUNCTIONAL DEMONSTRATION MANEUVER ABORT TECHNIQUE DEMONSTRATION

M = 6 - HIGH DYN PRESS. SAME AS 8 PLUS DEMONSTRATION OF REUSE

Note: Scheduled Maintenance/Ground Operation Provides Servicing/Maintainability Data GP74-1039-53


159



Approximately 45 instrumentation measurands are required as a minimum.

Thermocouples, pressure transducers and strain gages would be installed during

manufacture of the subsystems and calibrated after final experiment installation.

The instrumentation would be designed to be compatible with the airborne instru

mentation recorder and telemetry system in the X-24C.


160



. • STRAIN GAGES - 6 - VERIFY PREDICTED FIN OVERALL

LOADS AND DISTRIBUTIONS

,~ • ACCELEROMETER -1 {TRI-AX} - VERIFY LOAD FACTORS

AND VIBRATION ENVIRONMENT

• THERMOCOUPLES (FIN) - 20 - VERIFY STRUCTURAL TEMPERATURES - MAGNITUDE AND DISTRIBUTIONS

• THERMOCOUPLES {ACS} - 9 - VERIFY COOLANT TEMPERATURES, HYDROGEN TEMPERATURES

9 PRESSURE TRANSDUCERS - 9 - VERIFY SYSTEM PRESSURES, COOLANT AND HYDROGEN FLOW RATES

G P 74-1 039-54


161


FLIGHT TEST TASKS

The tasks in which the experiment contractor would be involved during Phase III are

listed on the chart. The first of these tasks would be pre-installation experiment

checkout to insure that all experiment hardware is received at the flight test location

ready to be installed. The experiment will then be installed in the X-24C. This

involves mating all experiment elements to the test aircraft. The experimental fin

will be spliced to the basic structure, subsystems will be installed and all required

interfaces installed and mated. The final assembly of all experiment elements will then

be checked. This will include functional testing of the cooling system, a check of the

failure detection system, checks of structural splices and instrumentation checkout.

Experiment contractor support will be provided during the test program on a "as required"

basis.


162


FLIGHT TEST TASKS

• PRE-INSTALLATION EXPERIMENT CHECKOUT

• INSTALL EXPERIMENT IN X-24C

• EXPER1MENT CHECKOUT • STRUCTURAL SPLICES

• ACTIVE COOLING SYSTEM

• FAILURE DETECTION SYSTEM

• INSTRUMENTATION

• PROVIDE SUPPORT AND ANALYSIS AS REQUIRED GP74-1039-105


163


BASIS FOR PHASE III COST

Again, assumptions must be made to establish a basis for costing Phase III.

These assumptions are presented on the chart.

164


BASIS FOR FLIGHT TEST COSTS

• FINAL INSTALLATION OF EXPERIMENT PACKAGE AND CHECKOUT AT EAFB

• GROUND TESTS AND DATA ANALYSIS PRIOR TO FIRST FLIGHT

• X-24C AIRCRAFT FLOWN AND MAINTAINED BY NASA

• TEST FACILITIES/RANGE SUPPORT PROVIDED BY NASA

• DATA COLLECTION BY X-24C RECORDER/TELEMETRY SYSTEM

• DATA REDUCTION BY NASA

• MINIMUM CONTRACTOR SUPPORT

• 9 FLIGHTS AT 3 PER MONTH RATE

• LH~ AND ASSOCIATED EQUIPMENT AVAILABLE AT EAFB

o NO REFURBISHMENT OF EXPERIMENT HARDWARE GP74-1039-51


165


FLIGHT TEST COST AND SCHEDULE SUMMARY

Duration of the flight test phase was assumed to be 7 months, concluding at the

end of 31 months from experiment ATP. The estimated total man-hours and cost are

9,420 hours and $263,000, respectively. It was assumed that support for the actual

flight test would average 720 man-hours per month for the three month period. This is

equivalent to 4 to 5 men and may be slightly low.

Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b), 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

166



PHASE m FLIGHT TEST COST AND SCHEDULE SUMMARY

ACTIVITY MONTHS FROM GO-AHEAD 1 2 3 4 5 6 7 8 9 10 1112 13 14 15 16 17 181920 21122 23 24 2526127 28129 30 31 MAN HO U RS COST*

ATP i~

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL

PROGRAM DOCUMENTATION ----------- ---- -J-- - -- - ---- - -- - ---- ---- ---- - -- - ----TASK 8 - FLIGHT TEST. ___________________ ..1._ __ _ ____ _.I.. h=~~~l!').

DELIVERY OF EXPERIMENT ________________________________________________ .~).

LH2 SYSTEM FINAL CHECKOUT

2,340

7,080

AND INSTALLATION IN X-24C _____ -r ----- ---- ----- ----- ---T ---- ---f- ---- --~). } FINAL EXPERIMENT CHECKOUL_ T -J + -+ -_J

1 __ ----- -11---- - ..1._ - ---- ~>-

FLIGHT TEST - 9 FLTS@3/MO------------r-l-r-----r------r--------- ________ '--'r!---'-"""'Ii""..... ~~~~ ~~ j ANAL VSIS AND EVALUATION------ +- ----- -__ J __ J -- -- ---- -- ---- ----- ---'--1- =r ,.I AVERAGet

PROGRAM COMPLETE. _______________ -- - ---- - ---- - J J -----f-I---- --- --- -J---M~

*Constant 1974 Dollars PHASE In TOTALS 9,420

65,000

198,000

97,000

(101,000)

263,000

GP74 1039·55 Information contained herein is privileged or confidential information of McDonnell Douglas Corporation and exempt from public disclosure under subsection (b). 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

MCDONNELL AIRCRAFT COMPANY .

167



The total active cooling system flight experiment cost and schedule are summarized

on this chart along with the total manhours involved. The program spans a total of

31 months, approximately 115,580 manhours would be required and the rough order of

magnitude cost is $5.8 million, given in constant 1974 dollars.

I nformation contained herein is privi leged or confidential information of McDonnell Oouglas Corporation and exempt from public disclosure under subsection (b), 5 USC 552. Do not disclose outside recipient organization of U.S. Government.

168




ACTIVITY MONTHS FROM GO-AHEAD 1234567 8 910111213141516171B19202122232425262728~3031 MANHOURS

ATP ~}

TASK 1 - PROGRAM MANAGEMENT, INTERFACE COORD AND CONTROL, PROG RAM DOCUM E NT A TI ON . ________ .I=*:~~~~~::::;::*=*=:=*=***=*=*=~~=:=~=**=**~

PHASE I - DESIGN, DEVELOPMENT AND TEST. _____________ . ____________________ I=*:~~~~~~~~~:;::;::;::;~~~""

PHASE II _ TOOLING, FABRICATION AND ASSEMBLY ___________________________ -r --- --- ~

PHASE ill - FLIGHT TEST________________ ____ _ ____________________________ ~~~~O'Jo

*Constant 1974 Dollars PROGRAM TOTALS

11,700

60,740

36,060

7,080

115,580

COST*

326,000

3,397,000

1,856,000

198,000

5,777,000 GP74 103960 Information contained herein is privileged

or confidential information of McDonnell Douglas Corporation and exempt from pub· lic disclosure under subsection (b). 5 USC 552. Do not disclose outside recipient organization of U.S. Government.


169


CRITICAL TECHNOLOGY AND LONG LEAD TIME ITEMS

Two major problems have been identified as falling within the category of critical

technology and/or long lead time items. The first is the development of the cooling

system coolant-to-hydrogen heat exchanger. Preventing coolant freezing while main

taining high heat transfer capability with low weight and volume requirements will

present a challenge. The second is attachment of the FDS sensing elements to the fin

surface material. Good thermal contact along the total length of the element will be

required, and element spacing will be critical. While not listed as a critical area on

the chart another area which would benefit from some immediate effort is definition of the

best type of abort maneuver. Our preliminary studies indicate high lift, low lift-

to-drag descents are excellent trajectory candidates. However, data is required to

define fin heating during high angle-of-attack operation.


170


CRITICAL TECHNOLOGY AND LONG LEAD TIME ITEMS

• ACTIVE COOLING SYSTEM COMPONENTS

. . • FAILURE DETECTION SYSTEM SENSING ELEMENT

ATTACHMENT

GP74-1039-58


171


X-24C PROGRAM IMPACT OF EXPERIMENT ACCOMMODATION

A number of items must be considered in the initial design of the aircraft if an

experiment of this type i"s to be accommodated. All interfaces and structural provisions

must be recogniz~d and defined. All special tool requirements must be defined and

included within the development plan. Aircraft weight and C.G. changes due to the

experiment package must be considered. A minimum cost experiment, and full flight

envelope capability requires full accommodation of the flight experiment in the basic

X-24C aircraft design.


172


X-24C PROGRAM IMPACT OF EXPERIMENT ACCOMMODATION /

• INTERFACES AND STRUCTURAL PROVISIONS

• FIN/FUSELAGE SPLICE

• UTILIZE BASIC SPEED BRAKE AND ACTUATION

• PAYLOAD BAY CONTAINMENT OF ACTIVE COOLING SYSTEM

• ACTIVE COOLING SYSTEM INTERFACES

• FAI LURE DETECTION SYSTEM INTERFACES

• INSTRUMENTATION

• SPECIAL TOOLS

• FIN/FUSELAGE SPLICE

• FIN/SPEED BRAKE ATTACHMENT

• CONTROLS, WIRING, PLUMBING, SUBSYSTEM MOUNTS AND SUPPORTS

• BASIC AIRCRAFT DESIGN

• WEIGHT AND C.G. ENVELOPE GP74-1039-57


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POTENTIAL ADDITIONAL FLIGHT EXPERIHENTS

The active cooling system flight experiment as configured for this experiment

definition study will stand on its own as an experiment package. It is not dependent upon

other aircraft subsystem (other than power) or experimental packages. However," it

could, with very little redesign and cost, integrate with an integral LH2 tank experiment

or with a LH2 fueled scramjet experiment. An additional experiment, with high potential

payoff would be flight demonstration of insulated actively cooled structure. This could

be accomplished with the active cooling system experiment package with the fin accommo

dating add-on surface insulation packets. These bond-on insulation packets could be

configured to cover the leading edge and the surface of the fin. With the fin, and

cooling system, designed for the uninsulated surface heat loads risk to fin structure

would be low. Some additional cost would be required for the development of the bond-on

packets.


174


POTENTIAL ADDITIONAL FLIGHT EXPERIMENTS

• INTEGRATE WITH INTEGRAL LH2 TANK EXPERIMENT

• INTEGRATE WITH SCRAMJET EXPERIMENT

• DEMONSTRATE INSULATED ACTIVELY COOLED STRUCTURE

GP74-1039-59

MCDONNEL.L. AIRCRAFT COMPANV

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nasa flight experiments x-24creport moc a3265 20 january 1975 nasa flight experiments x-24c final...

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