NASA Technical Memorandum 58256

{NASA-TH-58256) HIGH ALTITUDE AEBODYHAH,IC H84-16182

PLATFO:_H CONCEPT EV_LUATIOH &HD PROTOTYPE

_HGIH_ TESTING {HASA) 12q p HC AO6/MF &Of

CSCL 21E Unclas

G3/O "7 I I _88

B

High Altitude

Aerodynamic Platform

Concept Evaluation and

Prototype Engine Testing

J. w. Akkerman

4

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/

• January 1984

N/A

National Aeronautics and

Space Admin_stratfon
,, NASA Tachulcal Memorandum 58256

i'

t

HighAltitude

Aerodynamic Platform

" Co,.cspt Evaluation and

Prototype Engine Testing

J. W. Akkerman

_: Lyndou B. Johnson Space Center

Houston. Texas

i

i_. January 1984

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CONTENTS

Page i

ABSTRACT ................ . . ............. i

INTRODUCTION ............................. 1 "

CONCEPT ANALYSIS Oooooooooooooeooeomooooeeoe 2 _,

I

PPDTOTYPE POWER nSYSTEM DEFINITION ................... 13

ENGINE TESTING AND DEVELOPME_f .................... 27

,i

TEST SYSTEM ............................. 31

'' DISCUSSION OF RESULTS oooooooooeooooooooooooot 38

i OTHER OBSERVATIONS ......................... 50

:*_' CONCLUSIONS AND RECOMMENDATIONSnooooeooooaeooooooooo 52

REFERENCES .............................. 54

,_ APPENDIX A - COST SENSITIVITY ANALYSIS AND AERODYNAMIC PERFORMANCE i

__2 IMPNOVEMENTS FOR A HIGH ALTITUDE AERODYNAMIC PLATFORM I

C PRELIMINARY DESIGN CONCEPT i _:_. * .............. A-I

_ APPENDIX B - THE HAAP OPBRATING PROCEDURE t TTA-T- 2P558m ........ B-1

APPENDIX C - EVENT CHRONOLOGY _ oooo*ooeooooooeeoooo C--_

_ °

ii':

PRECEDING PAGE BLANK NOT FILMED

t

iii

¢
TABLES

Table Page

1 DESIGN FF.ATURES . . . . ................... . 7

2 LIFE CYCLE COSTS . . ..... . ................ 8

_q

I

FIGURES

I

Figure

1 Operational concept ...................... 3

2 Early vehicle concept ..................... 4

3 Vehicle concept 2 ....................... 5

4 Vehicle concept HAAP3 ..................... 6

5 Coat sensitivity ........................ 9

6 Winds at altitude ....................... ii

7 Solar challenger ........................ 12

8 Original power system schematic ................ 14

9 Engine side view ........................ 15

lO Engine top view . . . . .................... 16

11 Engine fo,_ and aft views ................... 17

12 Turbocharger detail ...................... 18 4

13 System schematic with 2-stage turbocharger ........... 19

14 System schematic with 3-stage turbocharger ........... 20

15 Engine assembly, front quarter view .............. 21

16 Engine assembly, front view .................. 22

17 Engine assembly, right view .......... . ....... 23

18 Engine assembly, left view .................. 24 ,

19 Stock and modified valve motion ................ 25

20 Piston cap ........................... 26

21 Oil circuit schematic ..................... 28

22 Stock engine performance .................... 29

23 Stock engine performance at altitude . , . ........... 30

24 Altitude test chamber ..................... 32

25 Instrumentation list . . . . . . . . . .... . ....... 33

26 Data program .......................... 34

27 Data printout ......................... 36

iv

(
Fisure Pase

28 DaCa sheet . . • . • • . ........... . ..... 37

29 Project ehronolosy summary .............. . . . 39

30 Clearance volume limits .................. 41

31 AitlCude/speed for various curbochar_ers .......... 43

32 Flow/pressure/temperature relationships .......... 44

33 Flow volume vs RPM .................... 46

36 Mass flow ......................... 47

I 35 Data sheet (44B) ...................... 49

36 Data sheet (predicted) ................... 51

]

V
ABSTRACT

A design concept has been developed for maintaining a 150-pound payload at

60,000 feet altitude for about 50 hours. A 600-pound liftoff weight aerodynamic

vehicle is used operating at speeds just sufficient to o,vercome prevailing winds

and powered by a turbocharged four-stroke-cycle gasoline-fueled engine.

Endurance of 100 hours or more appears to be feasible with hydrogen fuel and a .,

lighter payload.

A prototype engine has been tested to 40p000 feet simulated altitude.

Mismatch of "_e engine and the turbocharger system flow and problems with

,,. fuel/air mixture ratio control characteristics pzohibited operation beyond

40,000 feet. But the concept could reasonably be developed to function as

i analytically predicted.

INTRODUCTION

I"

I, High altitude flight continues to be of increasing interest among both

II_ civilian and military organizations. Most high altitude vehicles have been

i" developed for the purpose of transportation Also, there are needs for high

li_ altitude surveillance, communications relays, and certain scientific

!_ applications including atmospheric monitoring and earth resources. These later

needs have been met for the most part with transportation vehicles. Rockets and i_:!_

i_,_ balloons have been used in some cases.

Studies at NASA Edwards Flight Research Center in the mid-1970"s indicated

_i__'_I that the technology in low-speed airfoils and lightweight structures could be

_ used to fly at extremely high altitudes and at relatively low speeds. This

! technology has the potential of satisfying many of the high-altitude flight

_ requirements at relatively low cost. It offers a combination of high altitude

_ capability and maneuverability not available in transportation vehicles and

,,. reusability not feasible with balloons or rockets.

_ Previous work at JSC has supported a vehicle conceived at the Dryden Flight

• Research Center (DFRC) to take high altitude air samples (Mini Sniffer)

'_ (ref. 1). This vehicle supported a short-term mission and used a non-airbreathing engine. _ore recent applications are aimed at long-term operation

ii " supporting a communication relay station.

Various concepts have been proposed of airplanes and balloons in

_, combination with solar cells/batteries, remote microwave power, fuel cells,

_, solar cells/regenerative fuel cells, piston engine with various fuels, etc.

_: Studies have indicated that all of these ate technically feasiblle with varying

,_,_. req, irements for system development and operational characteristics.

_ The central issue at this point is the cost of the concept, both in the

development phase and its potential low cost in the operational phase. The

• solar concept (ref. 2), whether applied to a balloon or an airplane, awaits the

evolution of low-cost solar cells before it can become economically viable•

Also, for any significant duration, it will require development of an energy

storage system for nighttime operations. The microwave-energy-powered balloon

_ or airplane (ref. 3) will require the further development of a microwave power

. t_ansmission system and an antenna which is steerable to the degree required.

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Its operational cost would be strongly influenced by the efficiency o£ the

energy transfer system as well as its capital cost,

A primary issue is on-station flight duration. Some candidate concepts

could conceivably £1y indefinitely (microwave) and, therefore, minimize the

handling costs and hazards involved with frequent flight up and down through low

altitudes. However, ground turnaround operations and associated equipment will

still be necessary and will have to be accounted for.

An aircraft system, using two or three vehicles, could maintain a

continuous service with the convenience of easy access for maintenance of the

vehicle and the payload system (fi S. 1). "On-station" redundancy could be

maintained. The issue of low-altitude flight of this airplane could be managed

with a variable geometry wins, using a low performance/high strength

configuration through the turbulent atmosphere, and the high performance/lower

strength configuration on-station at high altitude. Studies at JSC have focused

on this approach.

The author wishes to acknowledge a_d thank Dr. Kraft for his decision t__

fund this activity. Also, the many individuals throughout JSC are to be

commended for their support during the testing, analysis, and data reduction.

The patience of management personnel in supporting an activity so far removed

from the areas of responsibility and expertise in which they are officially

involved is acknowledged.

CONCEPT ANALYSZS

The configuration concept in figures 2, 3, and 4 show the basis of the work

reported herein. The concept has a wing which potentially folds to form a

triangular truss to provide a higher strength configuration during low-level

flights. At about 40,000 feet altitude or above, where suet loads are reduced,

a winch system deploys the wins into the fully extended position for efficient

operation during the on-station cruise part o£ the flight mission. Figure 4

(HAAP 3) is for a fixed wing version for concept evaluation. The concept was

presented in November 1979 at a workshop session held at Langley Research Center

(LRC). Since then _C personnel have verified the estimates of structural

weight and aerodynamxc efficiency (ref. 4). They have suggested the addition of

a controllable flap on the trailing edge of the outboard wing section to provide

lift in the folded configuration. Recent JSC studies have focused on the system

engineering aspects to understand better what the primary cost drivers are.

Table 1 shows design features for each of four variations of the baseline

vehicle, plus a configuration conceived by Dr. Paul HcCready. The preliminary

cost sensitivity results are shown in table 2. The sensitivity study

(appendix A) was performed using a computer to calculate performance and total

program costs. As shown in table 1, a baseline has been established which is in

the middle of the range of estimates of performance for the primary elements --

engine, propeller, structure, and aerodynamics. This baseline was then changed

by 10 percent for each o£ the variables to assess their effects on the llfe

cycle costs. The results are shown in figure 5. As one might expect,

aerodynamic efficiency and weight are the major drivers. Also, the analysis

predictably shows that lower operational altitudes are less expensive, within

wind limitations and communications requirements.

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Studies of winds aloft (ref. 5) indicate that the choice of an operational

altitude is strongly influenced by winds. The results are shown graphically in

figure 6. Since the forward speed rcquired is a major influence on power and

energy r_quired, there will be an advantage in operating at the altitude of

"minimum wind," which varies with season and latitude. It might prove costeffective to operate with different aerodynamic configurations during different ,,

times of the year. All our studies have been based upon the assessment of

operations at 60,000 feet. However, operational costs would be lower if the :,

altitude could be reduced. Work should continue to improve the math model and

further refine the concept.

There are, however, definite limitations to the analysis. Specifically,

there is presently a serious shortage of aerodynamic performance data on

vehicles operating at low velocity at very high altitudes. These data will

_i likely not become available in the near future because of the lack of facilities

ii for this kind of testing. Another source of potential uncertainty is in weight

_ estinmting. The accuracy to which weights can be predicted varies with the •

I accuracy to which loads are known. In vehicles of the type involved here, where

I_ structural flexibility will interplay with the aerodynamic loads, prediction by

i_ analysis is difficult.

Ii_ These issues, along with the relatively low cost of the vehicle involved,

i_i_ indicate that the best progress might be made by moving from the purely

i;_, analytical approach for vehicle design into hardware testing in a real

_i_ _ operational environment. Early vehicles could use the payload capacity for

i_ data-gathering on well instrumented, lightly loaded vehicles. Each flight could

i!_ serve to further optimize the vehicle and its operations, leading to an early

definition of an operational configuration. Seemingly minor improvements can

r_ result in significantly lowLr costs as shown in the sensitivity cost delta

_ estimates in table 2 This kind of payoff could justify a fairly rigorous

effort for vehicle optimization even with a limited projected fleet size.

_ Because of this and the fact that an early demonstration seems important to

i_ some potential users, a flight test program probably should be the next step. A

i_ review of potential flight test vehicles indicates that McCready's solar powered

I_ airplane (see fig. 7) is suitable, within certain limitations, and could have

early availability. It is a lightweight vehicle with low power requirements i_

,,, The only serious compromise when compared to the "baseline" vehicle design is

i_* its frailty due to its fixed wing. Test operations would have to be limited to

ii_! fair weather. Dryden Flight Research Center (DFRC) is in the process of

_/_ constructing a similar vehicle, but metal is being used instead of the

!i_ lightweight plastics. This vehicle iS also a good candidate for demonstration

tests. Some lightweight sailplanes would also be acceptable. Any of these

,!iI vehicles can potentially be used to test any or all of the presently envisioned

power systems -- solar cell/battery, microwave, and piston engine. Also, any of

these power system alternatives can be configured to provide power for an

l electric drive propeller.

Early flight tests for evaluation of aerodynamics could be balloon-lofted

and released for simple glide testing, with the later addition of batteries and

a motor to extend the test time and raise the structural loads to operational

levels. As candidate power systems become available, they could be added _o

e2:tend the flight time and/or increase the payload carrying capacity. A test

system of this type should provide adequate concept verification and

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demonstration at a relatively low cost. Such a flight test program in a real

environment could conceivably provide a high fidelity test of a power system at

a much lower cost than a test cell under simulated conditions operating on the

ground. The operational handling experience and derived vehicle flight data

would be valuable **extra** benefits.

_J

PROTOTYPE POWER SYSTEM DEFINITION

:1

The early design concepts of figure 2 were discussed at the November 1979 4

meeting at LRC. The power system schematic is shown in figure 8. The early

design is shown iu figures 9, 10, and I1. The turbocharger assembly is shown in

five views in figure 12. The early configuration was tested using a stock OHC

model 200 engine and two Schweitzer turbochargers, a model 2IAq243 for the first

stage and a model $6-200 for the second stage (close to the engine). Both

turbochargers used identical turbines, the only difference being the choice of

housings for the two units. The housing selection is discussed in the testing

section.

These two turbochargers were later changed to use a unit by Rayjay for a

first stage and IHI as a second stage with much improved results. Figure 13

shove the original two stage concept schematically along with the

instrumentation points.

A second basic configuration was assembled using three stages of

turbochargers as shown schematically in figure 14. The all-up engine assembly

is shown in figures 15, 16, 17, and 18. Early tests were performed using the

stock configuration engine. Later, the cam was modified to give less valve

overlap and earlier exhaust valve opening (see fig. 19). The compression ratio

was also increased from about 6:1 to 11:1 using a pistolt cap attachment as shown

in figure 20. As discussed in the test section_ the final configuration was

reotricted to operation with reduced manifold pressure to avoid preignition

and/or detonation.

The interstage cooling between compressors consists of seven stainless

steel tubes of the bellows type flowing in parallel. The material was supplied

by the Anaconda Corporation and is normally used inside a braided steel wire

cover to provide structural support. In th.[s casej however, the pressure is low

• enough not to require external support. _".e first stage compressor discharges

into seven I-inch inside diameter tubes, the second stage into seven 3/4-inch

inside diameter tubes, and the third st,_e into seven l/2-inch inside diameter

" tubes. The first stage is cooled by ambient air in the test cell while the

other two bundles of seven tubes are covered with a large flex tub_ and water is

circulated between the tubes for cooling. In a flight configuration, the outer

shell would be omitted. Inside each of the seven air tubes, in each heat

exchanger, a thin strip of stainless steel (.020-inch thick) is twisted about

five times per foot. This causes spiral flow in the tubes to promote heat

transfer, Four of the spirals in each heat exchanger are left hand and three

right hand. The end fittings are fastened and sealed with epoxy. Silicone

rubber (RTV) is used to seal the endu onto the turbochargers. Set screws

provide mechanical attachment.

The innerconnecting fittings on the exhaust side of the turbocharger

serve as structure to hold the turbocharger package together and to serve as

13

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Ftgure 15. - Engine assembly, front quarter vtew

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';: Ftgure 16. - Engtneassembly,front, vtew

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_.. Ftgure 17. - Engins assembly, rtght view

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mounting. In each case, flanges were fabricated for the end connections, then a

pipe was fitted to connect the flanges as required.

The air aftercooler is fabricated of heat exchanger tube normally used to

vaporize liquid nitrogen. It was chemically etched from .095-inch thickness to

.030-inch thickness for weight reduction and to provide more flow area. It has ',

eight external radial fins 2 inches long, eight internal radial fins 1/4 inch

long, all on a basic l-inch tube. The material is 5052 T4 aluminum. Six of :,

these tubes, 5 feet long, are manifolded in parallel. The oil cooler/reservoir q

is fabricated of ten 0.025-inch thick sheets 2 feet wide and 4 feet long. They

are bent to maximize the air spacing and spaced 1/2 inch apart by an O-ring

sealed hoop 6 inches wide and 24 inches long. The assembly is clamped together

between two 3/8-inch plates with 20 number 10x32 bolts 6 inches long. The oil

pickup port for the engine oil pump is ducted to the middle of the bottom of the

reservoir with a 3/4 inch by .062-inch wall alumin_ tube. The oil returns from

the turbochargers, the valve covers and the crankcase vent with 3/A-inch hose to

the collector at the center of the top of the reservoir. The crankcase PCV

valve is connected to the intake manifold. Later tests used a separate oil

supply for the turbochargers and simply pressurized it to the first stage

compressor discharge pressure. The crankcase of the engine was l_ft connected

to the intake manifold. See the oil circuit schematic, figure 21.

The prototype test system included several variations in mechanical details

as will be discussed in the section on results. One major change from the

schematic of figure 14 was the relocation of the carburetor to include two

carburetors mounted directly on the cylinders. This and the shift to three

turbos were the only major changes made. An effort was made to increase the

displacement of the engine by increasing the stroke, thereby increasing the

engine flow, but little success was achieved because of mechanical problems.

No test results of that configuration were recorded. Space limits in the

crankcase required a pressed-together assembly, making the modified crankshaft

difficult to work on. Some success with camshaft variations caused the emphasis

to shift so that the approach involving increased displacement was set aside.

Other variations are mentioned in the section on "Other Observations.'* The

drawings for each of the mechanical details are on file if they are needed.

ENGINE TESTING AND DEVELOPMENT

The testing was divided into three phases: (1) stock engine, (2) stock

engine at altitude, and (3) prototype system tests. The results of phases I and

2 are shown in figures 22 and 23. Much experience was gained with the engine

during these tests. In phase 1, the performance was measured with various

ignition timing, lubricating oils, oil additives (molydisulphide), fuels

(aircraft and automotive), ignition primary voltage, exhaust conflgurations,

induction system configuration, etc., (tests A, B, C, D, E, and F of

appendix A). None of these items affected the performance significantly except

the ignition timing. (Nominal changes in the ignition timing consistently

degraded the performance slightly.) The reason that the engine is so

insensitive is not entirely clear, but is likely the result of a basically

efficient m_chanical design (as reflected by the insensitivity to lubricant

selection) and the low compression ratio. The most curious phenomenon is the

relatively good specific fuel consumption in spite of the low compression

27
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28

,. .... :_ ::.•....j
29
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]_.. ....::

....- ...................................; .......,......J
pressure (110 psi). The results of the tests o£ the stock engine were

consistent with published manufacturers" data.

The phase 2 tests were limited at low power (high altitude) by the

inability to reduce the load sufficiently. A smaller pump and a larger valve

were used in the load circuit in the phase 3 tests.

Phase 3 testing began with the assembly of the prototype engine with two

_d

Schweitzer turbochargere and continued with various modifications. These are

discussed ia the "Prototype Power System Definition" and "Discussion of 4

Results" sections of this report.

TEST SYSTEm _

The initial test system consisted primarily of a hydraulic pump mounted

directly Co the engine crankshaft and a flow restricting throttle valve to

control the load on the engine. A 1-foot long torque arm was mounted onto the

pump and a load cell was used to measure the force generated. An electronic

tachometer was used to measure engine speed. Oil temperature and cylinder

temperatures were measured with Chermocouples, and all the data were recorded on

a paper tape in engineering units. A strip chart kept a record of oil pressure

and engine speed. An automatic ignition cut off was used sensing low oil

pressure (below 10 psi) and/or excess engine speed (above 7,000 rpm).

This test system was moved to the vacuum chamber and reassembled to do the

initial altitude tests of the stock engine as recorded in tests C and H

(appendix A). Tbe test cell (vacuum chamber) is shown in the background of the

photograph of figure 24. The schematics of the test system developed over a i

period of time on an as-needed basis and are sho,_n in appendix B (Operating

Procedure). Some special test systems were used on occasion and are discussed

below in connection with the test results.

The data system is worthy of special mention. It required considerable

innovation because of the remoteness of the engine (in altitude chamber) and the

desire to be able to detect problems before damage occurred. The

instrumentation list is shown in figure 25 along with the readout mode. The

data logger proved to be a very convenient way of recording most of the data,

recording the data on a paper tape in engineering units. The data program form

is shown in figure 26 along with a typical data printout sheet in figure 27.

Three scrip ch_trts provided quick access to the turbocharger boost pressures and

the engine speed data. Six differential pressure gages were used to monitor

pressure drops in the four heat exchangers, the carburetor and overall across

the system from inlet to outlet (See the data sheet in figure 28). The speed

pickups were a problem on occasion. Phototransistorl_ were used in early

attempts, bvt then magnetic washers were added to the shafts of the

turbochargers to excite inductive pickups. Some problems were experienced with

the pulse/analog conversion, but the system worked satisfactorily most of the

time. The fuel consumption was measured with a calibrated volume in a sight

glass using a stopwatch to measure time to deplete a known volume of fuel.

The vacuum system in the initial phase was a single 350 cfm piston-type

pump. It was adequate for about 20,000 to 30,000 feet simulation when the

engine was running about 4,000 rpm. A second vacuum pump with 850 cfm capacity

31
Ftsure 24.- Altitude test chamber.

32 <

,2• 77 7 . .......................................... . r'j
1 Pressure Compressor (1) Inlet 0-30 PSIA .... (Data Loggerl

2 " " (2) " " (Strip

3 " " (3) " " Chart )

4 " Carburetor Inlet "

.5 " Intake Manifold

6 " Turbine (1) Inlet

7 " " (2) "

8 " " (3) " "

9 " " (3) Outlet "

10 " Oil (Guage) 0-100 PSIA

11 " Dyno Load (Guage) 0-2000 "

12 Torque, Engfne Output 0-50 Lb Force

13 APress. HX 1 0-30 " H20 (Manometers

14 " HX 2

15 " fIX 3 "

16 '* Carburetor

17 " Flow Meter

18 liX 4 "

19 Air Flow Carburetor Discharge 0-I00 CFM Air @ B PSIA "

20 Speed Eng. Engine (Mer.er) 0-I0,000 RPM

21 Turbo 1 Compressor (1) 0-200,000 RPM

22 " 2 " (2) *'

23 " 3 " ( 3) "

24 Temp. Cylinder Head, West 0-500°F

25 " " P East "

26 " Oil '*

27 " Compressor (I) Inlet -70-4200°F

28 " " (I) Outlet "

29 I* " (2) Inlet "

30 " " (2) Outlet " (Data Logger

31 " " (3) Inlet "

32 " " (3) Out let "

33 " Carburetor Inlet "

34 " " Outlet "

35 " Aftercooler Outlet

36 " Turbine (1) Inlet 500-2000°F

37 " " (2) " "

38 " " ( 3 ) " "

40 " Exhaust Cooler Outlet 0-200°F

41 " Cooling Water Inlet -60--+IO0°F

42 " " " Out let "

43 " Chamber Wall "

44 " Cooling Shroud -300--4100

45 " Cold Box Air Outlet (Meter) -60--+lO0°F

Figure 25.- Instrumentation llst.

33

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FLUKE 2240B PROGRAM FORM

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MQ'_IIOR CIIANNEL ,_INGLE GIIANNEL__ | DATE _.

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C_IA_INEL FUNCTION LIMIT A LIMIT B LIMIT lENSE VALUE AND

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} - T} Head Te1_p.Wout =

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5 T} Air Inl,_tTemp. •

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": 7 - T} Carbur.e,or Inlet

_L 8 T} Compres or #._ Inlet ... •

_f 9 - T} .lst Sta e Discharge i

'_.. I0 - T2 Exhaust Temp. lo

:-:__ II - T2 Turbine #2 Temp. .

-_: 12 - T2 12

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___ 14 "'T2 Turbine tt3 Temp. t.

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-_, P60Z05 16 - AO0 MV"'C)il;Pressure " J •

_. 17 - T2 Turbine Itl Temp. LIMIT B PROGRAM

_: 18 - T2 LIMIT SENSE VALUE ANO

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_ P6038_ 21 - _0 MV Load P_ess.

: ._ PllOgB 22 - _00 MV lorcue _m

-:- P601._3 2} - AO MV Comp]'essor #1 'In: et Press. _=

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_- P60127 26 - AO0 MV Comcessor 02 h let Press. =0

_,! P6014C 25 - 400 MV Cornfraser #3 In:eL Press. =_

::+ P6Ol z6.. ,oo,v CornL essor, P;;;,.

.:. 27 - A0 V ENg,-I PM #2 _a "_

i':" 2R - 4 V 4 VDC EL Inside CIamber _,

',::':;: 29 - a v 4 VDC _"Outs ................ ide _,,amber . ]

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:- _ 32 - 4 V #2 Speed RF _e

_ - _ tl lurbinl It_ Speed RF

P60605 34 .-400 MV Acel Bier Pre_s.

_18u=e 26.- Data program.

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4 7 _ _,C _ ___ Accumulator Temperature

4 _ "'' 'd' 'F Air Flow Meter Temperatuze

A4 _ _ _ Chamber Wall Temperature

4 _ - 4 _0 "_ Shroud (Cold Wall) Temperature

4 2 I 1 _5 ", Second Stage Discharge Temperature

41 ! 24,P _ ........ -Third Stage Discharge Temperature

t_ e _,_ _ Compressor 2 Inlet Temperataure

•_ . r, _, v Air Plow (CFM)

8 C,__ n V Exh. Pressure

7 1,1, c v Turbine 3 Pressure

_(, 1_21 nv_Turbine 2 Pressure

_ 1,7_ ,;_ Turbine 1 Pressure

_.4 2 _ 1 ,..v Accumulator Pressure

_' .__ F,._4C,, _ Turbine 3 Speed

• ? ¢,5! 4 ,. _' Turbine 2 Speed

!" ) ! f_ _ ",_, v _ Turbine 1 Speed

? o _j _ 1 _ V _ _xplosive Limit (Outside)

,_ 28 5rt 7_, V Z Explosive Limit (Inside)

'_ 2_ _.,',_C. _ Engine Speed (#2)

_ 2 _ ; _.,1_ . V Compressor 3 Discharge Pressure

'_- _ 1_9 _V -- _ompressor 3 Inlet Pressure

,_-"_ _ * _.(,1 ,. V Compressor 2 Inlet Pressure

;,_ ?_ (_.._C _v Compressor 1 Inlet Pressure

._., _? ._09 n.V Torque - Ft/Lb

._. 21 1,C,_e, n_ Load Pump Pressure

_. 17 > l_0_a _ • Turbine 1 Temperature

: _/_ I _ _',610 _._ Oil Pressure

_. ! _ 1(_[,_,. "_ Turbine 3 Temperature

_°. 1_ ,.0_ _ Free Air Temperature

_. I 1 I ?,7 _8 °_. Turbine 2 Temperature

._i_ 1 r _1),6 "_ Exhaust Temperature

_" 9 1 C.__,. _. Compressor 1 Discharge Temperature

¢_ _ Compressor 3 Inlet Temperature

:i/ _ 94,5 '_

' S _.._ _ _ Carburetor Temperature

_ -Compressor 1 Inlet Temperature

_, Cyli.nder Head Temperature (East)

": 2 1_ _? "_ Cylinder Head Temperature (West)

ii_' I :.,9__' V____ Oil Temperature

,._ "" - -- Engine Speed (#I)

' " f'_"t _':

Fisure 27.- Da_a printout.

• 36
37
was added along with a better exhaust cooling unit. Adequate flow was available

to keep the test cell purged with fresh inbleed air•

A unit was added for cooling and dehumidifying the inbleed air. This unit !

consisted of two aluminum 3- by 5-it cold plates attached to a spacer framework,

mounting the two cold plates about 6 inches apart. A series of aluminum strips

about 6 inches apart between the plates f,rced the air coming in at the top to

flow back and forth as it progressed through to the outlet at the bottom. The

entire unit was covered with foam insulation. Liquid nitrogen was used in the

cold plates • This unit kept the cold wall in the chamber from being covered I

over with frost so quickly and save better temperature control in the test cell•

The only real problem with the test system was the inconvenience of its

access after installation in the test cell. With seven fans for air

circulation, the test chamber was essentially "filled." Any work required _i

removal of several fans and, in most cases, the large aftercooler heat exchanger

on top of the engine.

The test procedure (appendix B) outlines the operation of the system and

the many safety features.

DISCUSSION OF RESULTS

This report presents almost 3 years of activity concerning demonstration of i

an operational assembly of hardware for powering a vehicle at 60,000 feet

altitude. The purpose is to aid those who may continue the effort to avoid the

same problems• A chronology of the major events is summarized in figure 29. A I

more detailed chronology is shown in appendix C, which includes 76 tests, 92

engineerin S orders, 57 purchases, 49 correspondences, six meetings, three j

analyses, and five drawings. Each is documented in the files and may be

reviewed by interested individuals.

The simplest summary of the results is as follows:

1. The goal of demonstrating 60,000 feet operation was not achieved - only

about 40,000 feet on two occasions•

2. While the particular combinations of hardware selected for test did not

function as desired, there seems to be no basic reason why the concept cannot be

made to w_rk.

The significance of the decision concerning compression ratio was revealed

early in the project. In the beginning, considerable effort was involved with

making the clearance volume in the engine automatically variable, to obtain the

best possible combination of efficiency and power output. However, as the

analysis of the overall concept was being completed, it became apparent that a

controllable clearance volume did not offer a significant benefit• It was

dropped for the time being, because other factors were more important in the

cost realm (see append!n A). The choice of a "fixed" clearance volume continues

to be a major factor in the engine design. It affects exhaust gas temperature,

overall efficiency turbocharger effectiveness, volumetric efficiency, and

ultimately the turbocharger compressor size. It also affects maximum allowable

manifold pressure which limits power output. These factors should be related to

38
ORIGINAL FA_ I_

OF POOR QUALITY

RTOPFORHAAPSYSTEHSTUDY..... > A

RANKIfiECYCLEANALYSIS M

I RANKINECYCLESYSTEMDESIGN J

J

' $50,000 FROMDIRECTOR'SDISCRETIONARYFUND _ A

S

OHCENGINEDELIVERY > 0

LANGLEYM_ETING ..... _> N

F

:_ FABRICATEBASEPLATE , ,_ M

A

!i It._P 2B VEHICLEDRANING ..... _> M

SEALEVELTESTSOFSTOCKENGINE , , > J

J

ALTITUDETESTSOF OMCENGINE(STOCK),. , _> A

"_,_ E&DPROJECTREVIEW . > S

_i_,_ NAAP-3DRAWING .... > 0N

I_- TESTANDC/O OF ALTITUDEENGINE "_ D

,_ FORTUNE500 PRESENTATION _

ii_, 2 TURBOSYSTEM(20,000 FT) SCHWEITZER _ M

'_' 2TURBO SYSTEM(40,000 FT) IHI ANDRAYJAY.... _ J

A

!. 2 TURBOSYSTEM(40,000 FT) IHI ANDRAYJAY , > S

O

,. SYSTEHREVIEW"- _ ' N

ii_ 3 TURBO-l,000 CCBUILDUP I O

i_ IHI, SCHWEITZER,RAYJAYTURBOSI J

=L F

..... 1,000 CC ENGINERODBEARINGFAILURE •- M

1,000 CC ENGINECRANKFAILURE • ._ A

70occENGIN*E2MOOCAM , > e2 J

BROKENVALVE .... >

!_: mo 3 CAM .... > S

ii:_. CAHSHAFT/TAPPETFAILURE,, _ 0

•, =CYCLING"FUEL/AIR MIX 0

,!', 2 CARRURETORTESTS...... m . i _ J

::,.- 83 F

.... M

A

Ftgure 29. - Project chronology summary.

t

39
C+}

one another and included in the math model outlined in appendix A for an

adequate understanding of the need for a variable compression ratio.

A natural initial choice is oi_eration at the smallest possible clearance

volume that will avoid detouatiou with the fuel being used and allow adequate

manifold pressure for sufficient climb power. But that requires early exhaust

valve opening to maintain exhaust gas temperature at reasonable levels for

sufficient turbine power. To get the best engine efficiency, the turbocharger

size selection is driven toward very small units to function with low

. temperature exhaust gas.

As in any power plant, the major consideration is the combustion process

which is primarily influenced by clearance volume. The combdstion should occur

in the best possible temperature_ pressure, and turbulence enviromnent to

produce the most energy release.

Theoretical minimum clearance volumes "-r the engine are shown in

figure 30. The actual clearance volume mini_ limit that can be used is

• primarily a function of the chemical octane of the fuel. But the temperature,

heat transfer, turbulence, ignition characteristics, basic combustion zone

=_ configuration, and fuel/air mixture characteristics also are influential. They

are sometimes referred to as "mechanical" octane of an engine. These factors

_- must be determined by testing. But after the limit is established, the :7--

'_- adiabatic compression temperature data (figure 30) can be used to make

_- appropriate adjustments for inlet air temperature limits. Also, if a fuel is

used with known different basic temperature limits, then adjustments in

:_ clearance volume can be made to run that new fuel. For example, the limit for

_ •

!-j_ 100-octane gasoline is about 1150 ° F • For 90 octane it is about 1,000 ° F .

In

'-_ operating at 550 ° R (100 ° ¥) the difference is 2.5 to 3.5 in. 3, a difference of

A0 percent. This dramatically affects exhaust valve opening timing requirements

and the resulting energy delivery to the crankshaft and to the turbocharger.

: As discussed in appendix A, efficiency is paramount in this concept. The

machinery must produce the most favorable situation possible fcr energy release

_ and then convert that energy into useful work. In this example, the energy in

the exhaust gas is used for air compression; therefore, the choice of exhaust

_ valve timing is involved. This affects the expansion ratio of the gases in the

engine and the relative amount of energy released to shaft work or to exhaust

/ gas _nergy. Thus, as the exhaust valve is opened earlier, the shaft work goes

down, !cwering the efficiency of the engine but increasing the exhaust gas

_= energy to power the compressors.

_ While the original plan for the project was to "match" the

engine/turbocharger combination, little was understood of the complex

: interrelationship of the many factors which affect the results. Much could be

done aualytically, but there are some fundamental problems. Turbochargers are

basically considered in the industry today as "volume" machines. Some of the

older companies in the industry still consider the flow primarily on a mass

_ basis because that relates well to the power involved in the machine. Of tours,

both mass flow and volume must be considered in any analyses, but the difference

is in the way of "thinking" about the situation. The volume basis relates well

to various engine displacements and is used in most data for users, particuarly

in compressor selection. The turbines are somewhat flexible in their operation,
4l
t!_

and typically companies will base their entire turbocharger "fmnily" around a

single turbine. A series of housings (nozzles) is made available so that the

user can power the compressor selected to whatever spee_ is needed to give the

pressure ratio desired. Therefore, in a company's "line," there will be a

particular combination of compressor and turbine housing that will give the

desired result fo_ a selected engine.

l In past experiences, no turbine/housing data have been made available--

only compressor data. The user selects a compressor to match the flow of the

engine, and "tries" various turbine housings (nozzles) to get the unit to

produce the boost in manifold pressure desired.

This situation is the primary reason that the project has extended so long.

The only option was to "cut and try." The first try was to use the units built

by Schweitzer. The results were very discouraging. Even with the smallest

housings available, little speed could be attained. The turbine itself was

_ basically too large - by at least a factor of 2 end very little boost was "

achieved. As the altitude was increased, the mass flow decreased, reducing the

speed and power in the turbine even more. See figure 31 (tests 3, 4, and 5).

t_ The next step was to look for the smallest possible turbine. The unit

selected was the IHI model 51B used on the 500 cc displacement Bonds motorcycle.

_,_i_ This violated . basic groundrule of using American built equipment, but there ,_: was no alternative. The Schweitzer unit was the smallest available in the

i_ United States.

ii_, Early tests with the IHI turbocharger combined with the Schweitzer unit

_ produced the results of tests 8, 10, and 11 (figure 31). With the additional

_ flow induced by the IHI unit, the Schweitzer compressor was flowing in a very

_!_ inefficient range, but the results were much better. Finally, a Rayjay unit was

_? selected to better fit the compressor flow situation and runs 18 and 33 produced

_ sea-level power output at 40,000 feet altitude.

;!/ Many "adjustments" were made in runs 19 through 32 attempting to increase

_,= the turbine speed and produce higher pressure ratios. Fuel mJ.xture ratio (lean

i!_ and rich) ohowed very little effect. Ignition timing diffrent from stock timing

ii_ had only adverse effects. Probably the most discouraging effect was the

_: changing of turbine housings. Smaller housings (nozzles) caused extra back

pressure on the engine and reduced the volume flow of the engine so much that

- the net result was a lower boost. Larger housings reduced backpressure, and

:_:' allowed higher volumetric efficiency of the engine, but the lower turbine speed

!!_ produced lower pressures in the manifold and less boost.

,:_: It became obvious at that point that something different was necessary.

_ The data showed (fig. 32) that increased mass flow save better performance.

',_ Also, higher compression pressure save higher exhaust temperatures. It became

apparent that all three units tested had best efficiency in the range of about

2/1 compression ratio in each stage. A pressure ratio in the range of 2.8/I,

_ ' needed for 1/2 atmospheric manifold pressure at 60,000 feet with two stages,

i required much more exhaust gas energy than three stages operating at 2/1.

It was decided that the system would work best if the engine flow could be

increased as well as changing to three turbochargers. A 1000-cc engine was

assembled, using s crankshaft with l/2-inch extra throw. The initial test

42

• _ . _._ _._ _ . -._ , . _ , .:.... : ..........._., .._,,-.................. .._ :- - . !!T.. _- -
K,- 9'

ORIGINAL PAGE I_

OF POOR QUALITY

'1

43
I,T I

ORIGINAL PAGE I$

OF POOR QUAM'P(_

2000- "

18

f_'_3 _ RUN"|

1I

1500"

U.

o

; EXhhUSTTEHPVS.

MANIFOLDPRESSURE 1000 HAAPENGINE

500,

o '_ lb i's z'o is ' "

MANIFPOLD RESSURE, LB/2INABS

6- PRESSURERATIOVS.

I! 33 11 18|

o= RUN# FLOWRATE

_ HAAPENGINE

_- 4 _ OVERSPEED

I--

_ 2

LOWERLIMIT OF OPERATION

i ! t , i

I i _ w w

0 2 4 6 8 I0

FLOWRATE, LBI MIN

Ftgure 32. - Flow/pressure/_emperatm'e relationships.

44

r
experienced a rod bearing seizure - probably because of inadequate oil flow. A

second test with larger oil holes, resulted in a "slip" of the pressed-together

crank _ssembly. These results were discouraging because the crankshaft appeared

to be the quickest and easiest way to significantly improve the flow. At that

point, it seemed reasonable that the extra flow needed might well be produced

with a camshaft modification to give essentially standard valve timing, but ',

increased lift.

The stock crankshaft yes reinstalled and tests continued, but with no

improvement in attitude capability. The boost was always good at high flow

rates but as the altitude was £ncreasedD the pressures decreased, the exhaust

temperatures decreased, the boost reduced, and the system would consistently

cease to function when the manifold pressure was below 14 psia. Finally, the

exhaust valve broke on one cylinder when the system's rpm and temperatures

became too high. It was believed that something must be done to increase the

flow (i.e., the power to the turbines) to get the boost desired.

At the outset, it was recognized that the clearance volume must be reduced

to get the best combustion efficiency at reduced manifold pressures. But that

is a risky thing to do on an engine with potentially high manifold pressure.

Reducing the clearance volume was postponed as long as possible in the hope that

an adequate demonstration could be achieved without the risk. It became

apparent that volumetric efficiencies of 60 percent typically would not suffice

=_ and that 80 percent was required. See the curves of figures 33 and 34. Another

_ problem with reducing the clearance volume was the reduction of exhaust gas

_ temperature because of the increased expansion ratio of the combustion products

,_: This increased fuel efficiency at the expense of exhaust gas energy. In this

,_= case, the expense was prohibitive. It became apparent that still another cam

=_ modification was in order to open the exhaust valve very early in the expansion

stroke. Also, with decreased clearance volume, little overlap of exhaust and

_ intake valves is required. The result was the choice of using a long duration

_ exhaust lobe, and to avoid excess overlap (because the operation would likely

_ require higher exhaust pressure than intake manifold pressure) use a short

duration intake lobe.

To circumvent problems with overpressuring the small clearance volume, a

pressure switch was installed in the intake manifold that would disconnect the

ignition system if the manifold pressure rose above 1 psig (15.7 psia). Initial

runs were very encouraging, showing volumetric efficiency at low speeds in

excess of 90 percent. But at higher speeds, it was only 50 percent. A

mechanical check revealed that the intake lobe was providing less than one-half

the lift it was built for. Modification of the intake lobe required a broader

tappet. The modified tappet had failed and worn the cam excessively and the

weld material used to reconfigure the cam separated. In any case, this caused

the low flow. The cam was reworked to the same configuration and testing

proceeded.

Surprisingly_ the flow continued to be low at high speeds. Volumetric

efficiency iu the range of 60 percent produced mass flow along the lower limit

of the established requirement (see fig. 32). The manifold pressure limit held

the mass flow even lower than in the u_odified runs where the manifold pressure

was allowed to go to 20 psia (see fig. 32). At this point probably the most

significant observation made in the project was apparent. At the low manifold

pressure imposed by the pressure switch, the engine performed unpredictably. It

45
10

0 I000 2000 3000 4000 SO00 6000 7000 8000 9000

RPH

Figure 33. - Flow volum vs RPH,

46

4_. _ _ . il _

• ; .
4'

ORIGINAL PAQIEI{i

OF POOR QUALITy

47

( J,

, .................................. : ........................... :........
seemed to have a mind of its own. On occasion it would run very well, only to

be followed by a period of extremely low-power operation - all on a 1- to

2-minute cycle. The ignition system seemed to be entirely functional. The

engine/mechanical system was verified to be "as designed." The fuel/air flow

system, however_ showed unexplained pressure-time variations that previously

were not considered carefully since moat testing was done at higher manifold

pressures. An exhaust gas analyzer was used to verify that the fuel and air

mixture was not being delivered to the cylinder with a consistent mixture ratio.

On more than one occasion, fuel was found puddling in various places - heat

, exchangers, accumulators, etc. This was disregarded because the system

performed predictably in moat cases (when the mass flow rate was high).

However, when the mass flow rate was reduced by lowering the manifold pressure,

the "puddles" apparently did not get "blown" along consistently, resulting in

periodic "off mixture" operation. Review of the data indicates that in every

case where the dynamic pressure of the flow is reduced, the system ceased to

: function properly. In early tests, this was caused by inadequate flow, giving '

inadequate turbine power. This was true, but not so much because of inadequate

flow as low "quality" flow. A few moments of extremely rich or extremely lean

fuel mixture would reduce turbine power enough to put the entire operation out

_ of commission. The only way to get the system up and going again was to 8o back

_ to sea level and start again. Consistently, whenever the manifold pressure got

below 1S psia, the system would "die.*' This became apparent when the pressure

-. switch was installed and all the operation was at the lower flow level. The

_ system could not be coaxed to work, even with the turbochargers disconnected.

'_*_ It would run properly with the stock manifold, however, so the problem was

narrowed to the aftercooler and the accumulators.

_L

_ As a result, the carburetor was moved downstream of the aftercooler but

-_ upstream of the accumulator. Two carburetors were used - one in each side of

the system. Operation was somewhat better, but cycling still occurred at a

_ somewhat higher frequency - a few seconds instead of minutes. This indicated

that the flow of fuel was possibly "hanging" on the walls of the induction

_ system - and as the amount of area was reduced, the amount of fuel and the time

_ involved were reduced. (The surface area was reduced by about one-half.)

Finally, the two carburetors were moved to the _ylinder _nterface downstream of

all the plumbing. This gave steady engine operation at a reasonable specific

fuel consumption. At last the engine would run steadily with less than one

atmosphere in the manifold as it had in the tests G and H two years before.

_: The only problem was further flow reduction. With the carburetor upstream

,_ of the accumulators, the flow through the carburetors is essential]y steady,

i_ allowing about four times the time for the flow to occu_ as in the case of the

cylinder-mounted carburetor. Using the effective fl¢_ area of the carburetor

L_" throat, the volume that was required to 8o into the cylinder and the time

available at 6000 rpm, it is clear that the flow in the carburetor throat was

very close to sonic in the cylinder-mounted :age.

At this point, part of the test facility being used was needed for another

test project. The engine project was unfortunately terminated.

In event that someone will pickup the effort and continue, the following

section, "Other Observations," outlines several of the "fringe" problems

encountered and the solutions to them. Figure 35 shows a set of run data (run

no. 44B). An analytically-derived set of operating conditions that should be

48
4

9

. ........ ........ .....j
demonstrable using the proper components is shown in figure 36. All the other

run data are available if needed.

OTHEROBSERVATIONS

Several problems came about which were not entirely related to matching the

turbochargers and the engine, mtt two deserving some mention are oil control and

cool ins.

}+

The problems experienced with the oil were of three types:

I. Cavitation of the pump due to low feed pressure and high speed

2. Air/oil separation (foaming)

3. Inadequate return flow to the smnp

i Inin some the OcMCaseengine s, all begins three ma to ycahav vitate e beenat occtheurrisuction ng at atthe about same 4timpe.sia Twhen he oilusingpump

17 SAE 10W oil at 250 ° F operating at 5000 rpm using the 5/8-inch I.D. lO-inch long

I_it suction tube as in the configuration. This was observed in the initial runs up

to 40,000 feet with the crankcase vented. Later, in an effort to use oil

if= cooling for the cylinder heads, an additional 3/8-inch I.D. suction line lowered

ii_ this to about 3 psia. In most tests, the crankcase was connected to the

_+'_ discharge of the first compressor, keeping the crankcase pressure at or below

++_ the pressure in the compressors and turbines. This kep _ the oil from leaking

"._ past the compressor seals. However, there is evidence that air and exhaust

leaked from the compressor into the crankcase cavity.

+:_ Oil consumption was a consistent problem during the entire altitude testing

:_ part of the project. In reflecting on the data, and after _olving items 2 and 3

,+,+ (air/oil separation and low return flow), it appears that the problem was with

++ the crankcase vent. Mechanically it is arranged on the camshaft drive gear with

i!!i+ the vented air passing inward through three radial holes in the cam sear and out

i_!i _ through a passage cast in the crankcase adjacent to the cam bearing. The unit

_+!' is designed to handle normal ring leakage, but it is possible that compressor

,+ and turbine seal leaks go for three turbochargers was too much for this unit to

,:/ handle and caused oil droplets to be blown through the vent mechanism. Future

testing should include a clear line (tyson ,r glass) to verify this as a

_, problem. The vent could be connected to an external supplementary tank through

_i+ a relief valve (to maintain sufficient crankcase pressure to avoid pump

!i_ cavitation, and stay below first-stage discharge pressure) to verify leakage.

Removal cf the turbochargers from the crankcase oil system to a separate supply

i: eliminated the oil loss and verifies this hypothesis.

,I

.... Power loss in early tests might have been caused by excess oil in the

crankcase, but carburetor location and poor fuel mixture ratio control was the

+L probable cause. The initial fix was to provide an air vent from the vapor space

in the sump to the top of the cxankcase. This allowed the oil to return to the

sump along with whatever bubbles were entrained without the vapor going upstream

in the return oil flow passage. The return oil flow passage from the crankcase

to the unit is about 1 inch in diameter. It was surmised that the 5-gallon

per minute oil flow was too slow if the gas from the accumulated bubbles was

5O
I

,., OF POOR QUALITY

oo, L--_I,

"r?l _ " _ _

• _

t _

• !

• c_

.51
passing back upstream into the crankcase causing overfilling in the crankcase

and a large power loss. The addition of the vent line seemed to fix the bogging

problem at that time, but no specific teats were run to verify it. All later

power loss was accounted for by lack of turbocharger performance.

Later, there continued to be probIems with partial loss of oil pressure

after Ions running - 10 minutes or more. The cause was surmised to be the long

flow path in the sump oil cooler, allowing an accumulation of bubbles to become

entrained with the oil. Thirteen 2-inch diameter holes were made in each of the

10 fins extending through the oil tank, and no further problems with oil

pressure degradation were experienced.

In future testing, consideration might be given to using a separate oil

pump operating at a very low speed to avoid cavitation at even I psia. This

would allow venting of the crankcase to ambient air and would avoid reingestinS

the ring leakage and the exhaust leakage past the turbine seals into the intake

system.

, The other area of considerable consternation was cooling, or more generally

thermal control. In some instances low temperature was a problem. This area

will be greatly simplified when the environment is not artificial and when the

power system behaves as designed. The overheat situations were always

_: associated with high power tests to maintain sufficient flow for the desired

":_! bloost ikewisepressures. cases withThelowlowpower temperature (idling) siituat n iaonswait mthat ode gave to achieve problems lowwere cell

temperature.

_ A notable exception is the problem with low oil temperature. This is the

result of excessive fin area. The design was intended to be that way, with the

_ plan to remove fin length as required when the system was functional. This low :

_ temperature did aggrevate pump cavitation, however, until much of the fin area

_ was covered.

_ The addition of air intercoolers between the compressors appears to be an

!_ attractive feature. All the best turbocharser performance was recorded with

cold air. This is a significant efficiency consideration in that the

_o compression becomes closer to a constant temperature process. Also, with the

_. denser air, the compressor gives higher pressure ratios with lower turbine speed

•_ requirements. The lowest temperature of operation was on run no. 18 at -8 ° F in

_" the test cell. Operation at -70 ° F can be expected to be much improved -

especially if effective intercoolers can be used between each stage. The

_ bellows type tubing with a spiral turbulator, as seen in figure 16, yields an

_i effective intercooler with low pressure loss. Mechanical vibration and fatigue

_ cracking may be a problem with this design, however, so some life testing needs

to be done.

CONCLUSIONS AND RECOMMENDATIONS

The only conclusion that can be reached here is that matching engines and

turbochargers in the small size range, where only limited variations of sizes

are available, is difficult. Limited published data on flow characteristics of

the various elements makes it difficult to do analytically, thus requirin S

52

• &
II 

' experimentation. However, there seems to be no reason why the concept will

not work.

After the concept is basically demonstrated, there will be a great need

I,

to establish the altitude limit of operation of the concept as it relates to

the specific fuel consumption. Specifically, with reduced clearance volume (to

the limit of power capacity for climb), the question becomes: "How early must

the exhaust valve be opened to provide sufficient exhaust energy to run the

curbochargers?" At higher altitude, more of the fuel energy must be used for

gathering air, thus requiring a higher specific fuel consumption, This effect

will likely begin to be a significant factor at about 50,000 feet. Above that

altitude, the specific fuel consumption will go beyond the typical 0.5 pounds

per horsepower hour feasible at the lower al_itudes.

It is recommended that the project be continued with basically the identical

!! configuration except with improved fuel control, possibly with a fuel injection

! system located at the cylinder inlet. i

• IiI,

It

'1

1!,

53

.... i .........1!..................................... : r,i ' ........ • " ..... ........ .......... ' -
REFERENCES

1. Reed, Dale R.: High Flying Mini-Sniffer RPV: Mars Bound7 Astronautics and

Aeronautics, vol. 16, no. 6, June 1978, p 26.

2. Phillips, William H.: Some Design Considerations for Solar Powered

Aircraft. NASA Technical Paper no. 1675, June 1980. d

I

3. Turriziane, R. Victor: Sensitivity Study for A Remotely Piloted Microwave

Powered Sailplane Used as a High Altitude Obeservation Platform. NASA

Contractor Report 159089, June 1979.

i

4. Morris, C. E. K., Jr.; and Liu, Grace C.: Performance Evaluation of a High

Altitude Airplane Platform with a Cryogenically Fueled Engine. NASA

TM 83184, Dec. 1981.

5. Strganac, Thomas W.: Winds Study for High Altitude Platform Design.

NASA Wallops Flight FacilltyRP-1044, Dec. 1979. 12

>

- )

' _,

,!

54
®,

i_t ,t

\' i

i'

APPENDIXA

l,

I;

!,.

Li COST SENSITIVITY ANALYSIS AND AERODYNAf'IICPERFORMANCE

I; IMPROVE_IENTS FOR A HIGH ALTITUDE AERODYNAMIC PLATFORM

li_i PRELIMINARY DESIGN CONCEPT

;L'

,I Robert Zehentner

"I"I Pro§ram Development Office

.I

I

t

i_

v

a

:: Lyndon B. Johnson Space Center

Houston, Texas

ii

,: March 17, lg81

z
ii CONTENTS

ABSTRACT ............................... A-1

INTRODUCTION ............................. A-1

COST SENSITIVITY ANALYSIS ....................... A-2

Results . . ........................... A-3

AERODYNAMIC DESIGN IMPROVEMENTS .................... A-7

Airfoil Selection ........................ A-7

Aircraft Parasite Drag Reduction .................. A-8

Induced Drag an, W_ng Optimization Analysis ........... A-]O

Effect of Aero Efficiency Improvements on Cost Effectiveness . . . A-II

RECOMMENDATIONS ............................ A-12

I REFERENCES .............................. A-13

_ BIBLIOGRAPHY A ]4

,I!L_ ooooooeeiooeeeooeooooooooeee°

:?

3

_!:i!'

,L

3 !i_'

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!._

1

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TABLES

Table Page

1 Aerodynamlcs/Performance ................... A-16

2 Cost Sensitivity ....................... A-17

3 Design Features ....................... A-18

FIGURES

Figure

1 HAAP "-__oneepu ......................... A-19 •

2 Cost Effectiveness vs. Parameters ............... A-20

' 3 Cost Sensitivities A 23

_" 4 Cost Effectiveness vs. Usefulness Factor ............ A-24

5 Cost Effectiveness vs. Fuel/Payload Ratio .......... A-24

...... 6 Range Optimization of Lift Coefficient ............ A-25

7 Endurance Optimization of Lift Coefficient .......... A-26

, 8 Wing Optimization - Climb Configuration ........... A-27

A- iv
®

ABSTRACT

A cost sensitivity analysts was conductedon a remotely-piloted Htgh

Altttude AerodynamicPlatfom (lt_P) preliminary des|gn concept. Ustng a stngle

mtss|on proftle math model, all cost-affecting parameters were Independently

varied from a defined reference vehlcle. The cost sensitivity of each performance, operations, and financing parwneter was established by calculating the

veh|cle cost effect|veness changefor + 10 percent variation from the reference

value of that parameter.

The results of the sensitivity study suggestedfurther analysts to tmprove

vehtcle 11ft anddrag coefficients. A htgher performanceatrfot1 and vehtcle

drag reduction resulted tn greatly tmprovedvehtcle cost effectiveness as

expected, _h_le c11mbhorsepowerhad very 11ttle Influence on cost effectiveness.

INTRODUCTION

Geosynchrenousearth orbtt satellites have proven their effectiveness as

communicationrelay systems for large area coverage. They are less costeffective for smll areas wtth relatively few users and for mobtle receivers

where antenna stze ts 11mtted. More efficient 11mtted area coverage maybe

accomplishedby stationing of a Htgh Altttude AerodynamicPlatform (HAAP)

at an altttude wherewtnd gust frequency and velocities are mtntmal and transmitter coverageadequate. _ny posstble communicationrelay applications

tnclude ne_s medta onstte coverage, county and state law enforcement stattons

and vehicles, and low usage telephone systems. i

Ptloted vehtcles requtre 11mtted f11ght duratton resulting tn an Increased

frequency of f11ghts for continuous coverage and higher operations costs. A

remotely-piloted vehicle does not haveth|s constraint and makeslower veMcle

costs posstble by eliminating ptlot wetght and the need for cockpit environment.

The cost of a remotely-piloted HAAPmust be competitive to replace exJsttng

communicationrelay systemsand for other uses. Thts report describes a cost

_ens|ttvtty analysts of a nomtnal HAAP.

A-1
COST SENSITIVI[Y ANALYSIS

The cost sensitivity analysis was performed on the HAAP preliminary "

design concept originating from the Systems Design Office in late 1979 :

(references l and 2). An uvaluation of the vehicle concept was performed by

Langley Research Center (ref. 3) which verified the feasibility of the HAAP

project and suggested design improvements in airfoil selection and wing optimi- i

zation. Table l compares the preliminary design aerodynamic parameters with

evaluations of the same made by the Langley study and the Entry Aerodynamic

Design Section at JSC (ref. 4). The consistency of parameter values is an

indication of the validity of the original design concept. The last column in

the table shows the values selected for the nominal vehicle used as a reference

in the sensitivity analysis.

The configuration analyzed is illustrated in figure I. It features out- %

board wing sections that are folded during takeoff and low altitude climb for

reduced sensitivity to atmospheric turbulence. The outboard . :ctions are

deployed to the cruise configuration on reaching the altitude ._twhich the full

"_ wing area is required for continued climb.

_ The first step in conducting the analysis was the design of an analytical

-' model simulating a single HAAP mission profile: climb to 40,000 feet, deployment

of the folded outboard section, and climb to 60,000 feet where the mission is

completed in cruise as the fuel supply is depleted. The math model includes

all mission parameters affecting flight cost and hence vehicle cost effectivei ness. Aerodynamic and engine performance parameters examined were cruise lift

---e and drag coefficients, propeller and drive train efficiencies, wing area, vehicle

dry weight, engine power available and specific fuel consumption. Operations

and mission profile parameters included in the model were cruise altitude, fuel

'

_ and payload weights, operating costs per flight for man-assisted launches and

}

i facility cost per hour for hangar storage. A computer program coupled the

mission profile math model to vehicle lifetime and cost, percentage of vehicle

life spent in cruise, and an overhead cost factor by which all cost inputs are

multiplied, lhe program calculated vehicle cost effectiveness in dollars per

pound of paylo_d per hour of cruise at 60,000 feet ($/IbpL-hr). Other outputs

A-2
of the program includea cost distributionbreakdownfor vehicle,operations,

fuel, and facilityas percentagesof totalmission cost, alongwith wing

loading,fuel-to-payload-weightratio,and usefulnessfactor (definedas the

ratioof fuel plus payloadweight to total takeoffweight). These outputs

were desiredto help simplifyrecognitionof cost effectivenesspatternsamong

combinationsof originalinput parameters.

} If any furtheranalysesare carriedout, some minor refinementsin the i

computerprogramare suggested. Outputs of total singlemission cost and total

h,

i lifetimecost would be helpful in evaluatingresults. Also, the analytic

! missionmodel assumeda static rate of climb definitionwhere all excess power

i,"

i availableover power requiredwent into raisingthe vehicle,assumingthe

_, aircraftheld a constantvelocity. A dynamicrate of climb definitionwould

}!_ slightlydecreaserate of climb performanceand would be more appropriate.

!'_ Climb power available,however,had littleeffect on cost effectiveness;therei_i fore the validityof the sensitivityanalysiswould be unchangedas long as

ii_ overallconceptdoes not change a great deal.

,_j_ To determinecost effectivenesssensitivityto individualparameter

_, changes,a referencevehicle was first defined. The analysiswas con-

!

_ ductedby varyinga singlecost-affectingparameterwhile holdingall other

parametersat their referencevehiclevaluesto determinethe independentcost

i_: effect. Mission profilecomputerrunswere made with parametervaluesboth

above and below the referencevaluesto generatecost effectivenessvs. para-

_ meter curves for each parameter. These plots are displayedin figure2.

_:,_ Results

_,/' Figure3 shows the relativeindependenteffectsof variousparameterson

referencevehiclecost effectiveness.The referencevehiclehad a cost

!:, effectivenessvalue of $.103 per pound of payload-houron-station. This value

_'J is based on a vehicleof nominalperformancepa,.ameterswith a 5-yearvehicle

_' lifetimecost of $50,000,an operationscost of $150 per flight,a facility

cost of $.25 per hour of vehicleon-stationtime, and fuel cost of $30 per

flight. The $50,000aircraftcostwas an estimatebased on the cost of similar

vehicles. A 50 percentutilizationfactorgives an on-stationlifetimeof 2.5

years at 60,000 feetwith the other half of its life spent largelyin the hangar,

since climb is only 6.5 percent of total time aloft. With low turbulence

A-3
+It)

and gusts in the on-station cruise configuration, a 5-year lifetime seems

reasonable. The $1bO operations cost per flight represents an estimated

three man-hours for preparation, launch, and recovery at $50 an hour. The

$.25 facility cost per hour on-station is equivalent to $180 monthly rent for

a hangar shared by two vehicles. Fuel cost of $30 per flight is based on 150

pounds of gasoline at $.20 per pound. An overhead factor of 2.0 doubles all

cost parameters to account for financing and insurance, and results in a cost

per flight of $535.90 and a total lifetime cost of $338,000 per airplane. The

cost distribution is: operations 56.0%, vehlcle cost 29.6,%, facility cost 3.2%,

and fuel cost !1.2% of total cost before inclusion of the overhead factor.

From figure 3, it is seen that reference vehicle cost effectiveness is

more sensitive to the time-weighted average lift coefficient than any other

parameter. S, nce the vehicle spends approximately 95 percent of the reference

mission profile in the cruise configuration, the time-weighted average effectively

becomes the cruise lift coefficient. The figure displays the change in cost

effectiveness for I0 percent variations from the reference value for each of

the parameters. A I0 percent change in cruise lift coefficient (0. I) yielded

a cost effectiveness difference of $.009 per pound payload hour, or a 9.23 percent change in the base vehicle cost effectiveness of $.103. Table 2 gives

the dollar _,alue of this I0 percent parameter aesign improvement be,sed on one

reference vehicle lifetime as already described. In practical terms, then,

a I0 percent design improvement in cruise lift coefficient (from CL : 1.0 to

I.I) on the reference vehicle would yield a savings of $31,200 or 9.23 p_,rcent

of the total lifetime system cost of $338,000. Over the lifetime of one hundred

HAAPvehicles, the design improvement would save $3.12 million, well worth using

the most efficient airfoil section design available. This great improvement

in cost effectiveness is largely due to the large increase in endurance time onstation.

Lift coefficient is followed by cruise altitude in cost sensitivity. A

I0 percent cruise altitude change yields an 8.75 percent change in cost effectiveness. The cruise altitude parameter does, however, have limitations of

smaller area coverage and greater high gust frequency at lower altitudes, but

these problems are minimal at the 54,000 foot level, a I0 percent decrease in

the reference vehicle cruise altitude of 60,000 feet. Single vehicle lifetime

A-4

.... . , . : •
savingswould amountt_ $29,600,largelyresultingfrom an extendedcruise

durationmade possiblewith additionalfuel not used in the last6,000-foot

climb increment. Fuel saved is substantialbecausethe climbmode becomes

relativelyinefficientas the aircraftceilingis approached.

Loweringthe referencevehicleempty weight IW 10 percent (54 pounds)

improvescost effectivenessby 8.26 percent for a vehiclelifetimesavings

of $27,900.

It must be rememberedthat the sensitivityanalysisis restrictedto a

parameterrangewithin_ lO percentof the referencevehicle base value. Outside that range,sensitivitieswould not be the same, so that a largervariation

in parametervalue would not necessarilyhave proportionatelythe same effect

on cost effectiveness. It shouldalso be noted that the lO percentvariation

_ used in this analysisis arbitraryand may not representthe probableranges

8,

_ of parameters,which are discussedbelow.

4 Usefulnessfactor,definedas

FUEL WEIGHT+ PAYLOADWEIGHT 

U.F.= 

TOTAL WEIGHT

-q" '1

"_ is plottedvs. cost effectivenessin figure4. Cost effectivenessimproves =

rapidlyas usefulnessfactorincreasesup to about 0.5, but very little

- thereafter. This suggeststhat fuel, fuel tank, an_ payloadshouldcomprise

at least 50 percent of the total takeoff weight to achieve the most cost-

* effectiveflights. Figure5, fuel weight/payloadweight ratio vs. cost

_o effectiveness,showsminimum cost when fuel and payloadweights are about

_ equal. Therefore,a generalrule of thumb for best cost-effectivenessis

to plan for fuel and payloadeach to represent25 percentof the total weight.

A final observation,noted throughoutthe analysisprocess,is that climb

_ perfomance ha_ negligibleeffect on cost effectiveness.This is true if the

40,000foot wing deploymentaltitudeand 60,OQOfoot cruisealtitudeare both

attainable. As longas these altitudescan be reached,climb lift and draq

coefficients,time to climb,an lfuel used in climb have littleeffect on the

totalmission since climb time is generallyonly 5 to 10 percent of time aloft.

For this reason,design improvementeffortsdiscussedin the secondpart of

this reportwere concentratedon the vehiclecruiseconfigurationwith later

checkingback for adverseeffectscaused by the climb configurationperforli_ance.

&-5
!

To establisha crediblebase for comparisonwith other typesof limited

area c_unication relay systems,a realisticcost effectivenessband _f

performanceshouldbe defined. By estimatingthe most optlillisticand pessimistic va]uesfor each oi'the cost-affectingparameters,a worst to best case

vehiclecost effectivenessrange is detenlIined.[able 3 giw._ the parameter

valuesfor the worst and best case vehiclesas well as for the reference

vehicle. The cost effectivenessvalue rangesfrom $.66_ per pound payloadhour for the worst case vehicleto $.021 per'pound payload-hourfor the best

case vehicle. The referencevehicle has a cost effectivenessoi $.103 per

pound payload-hour,much closer to the optimisticperformancevehiclethan

the pessimistic. With this cost effectivenessrange definition,a comparison

of the HAAP vehiclewith presentmethods of communicationrelay including

shortwave radio,helicopters,hot-air balloons,and geosynchronoussatellites

could later be performed.

To get a feel for the upper perfoniiancelimitsof a HAAP systemderivative,

a specialhigh altitude,high performancevehiclewas evaluatedfor feasibility

with littleregardto cost effectiveness.Performanceparametersare given

in table 3 and show a vehicletakeoffweight of 300 poundscarryingII0 pounds

of hydrogenfuel and a 5 pound payloadto a 90,000 foot altitude,where cruise

lasts for slightlyover a week (175hours). The cost was $2.80 per pound

payload-hour,over four times that of the worst case performancevehicle.

A-6

: :_mml Ir . -
AERODYNAMICDESIGN IMPROVEMENTS

The cost sensitivity analysis illustrated those areas where design

improvement would be most productive in iml)roving IIAAP cost effectiwness.

As a result, ._rodynamicdesign efforts were concentrated in the vehicle

cruise configuration.

Alrfoil Selection

An airfoil design for use on the HAAP vehicle must satisfy many requirements for efficient flight through its greatly varied regime. It must main-

,I rain effective perfori,_ancethrough a Reynolds number range of 500,000 to

I

!:. 2,000,000. Since extended cruise duration is highly desirable, the endurance

i,I ratio, CL3/2/CD' should be maximized. Low parasite drag at high angles of

,_ attack is also necessary. To achieve these characteristics, the airfoil

_ must have minimum skin friction drag and flow separation on the airfoil upper

,, surface Separation results fr(_11adverse pressure gradient regions where the

upper surface flow decelerates to achieve pressure recovery. Laminar (low

.i::!,,., skin friction) flows have minimal energy,to maintain attachment through these

regions. State-of-the-art airfoil technology deals with this problem by con-

" i_ trolling the airfoil velocity profile and, hence, pressure distribution. The

_ upper surfac_ pressure distribution desired is relatively flat (constant velocity)

il-, back to about 40 percent of chord, where controlled transition from laminar to

_ turbulent flow takes place and an abrupt pressure rise (adverse pressure gradient)

_L occurs. The increasing pressure region develops a conti_uously decreasing

p_essure gradient. This has a favorable effect on the turbulent boundary layer

" development nece_._ry for pressure recovery. The adverse pressure gradient

reduces as the trailing edge is approached. For best performance and highest

i,i lift, the optimal pressure distribution is one in which a nearly separated

turbulent boundary layer exists everywhere behind flow transition from laminar

'I to turbulent on the upper surface. The lower surface profile should be nearly

,_' flat, producing a gradual flow acceleration once past the leading edge, Laminar

flow is easily maintained since th_ pressure gradient is favorable. In summary,

a high 'performanceairfoil for the IIAAPvehicle should have a pressure distribution described by a laminar region back to about 40 l)ercentof chord where

I

i transition occurs and where a turbulent boundary develops from the rapidly

"_

decreasing adverse pressu_'egradier_t.

A-7
Of all the airfoilsconsidered,two fit the above descriptionexceptionally well. The WortmannFX7_-CLS-14G(reference5) and the LiebeckLIOO3M

(reference6) had comparablemaximum liftcoefficientsof 2.2 at a Reynolds

number (Re)= I X lO6. The Liebeckairfoilwas selectedover the Wortmann

for severalreasons. First, the LIOO3Mwas designedfor high aspectratio

applicationand so has a highermaximum thickness-to-chordratioof 18 percent. This enablesuse of a more substantialspar to handlespan-wisewing

loadings. Secondly,the Wortmannairfoilwas optimizedfor high lift and large

glide (CL/CD) ratioswhereas the Liebeckwas developedfor long durationsailplane applicationswith an emphasison high endurance(CL3/2/D)C ratios. The

designcriteriafor both airfoilsevidencethemselvesin their drag polars.

For the Liebeck LIOO3M,CD remainsabout constantup to the stall angle of

attack (ref.6, figure 12) whereas the drag polar for the FL?4-CL5-140(ref. 5,

figure5) shows a large increasein CD at higherCL'S. Therefore,the LIOO3M

airfoilwill give better cruisedurationfor the HAAP vehicle.

Designconditionsfor the LiebeckLIOO3Mat Re = l.O X lO6 includeCL = 1.8

and CD = 0.009 at an If.2° angle of attack. These valuesgive an airfoil

_ glide ratio (CL/CD) of 200 and an enduranceratio (CL3/2/D)C of 268 -- ideal

for the HAAP vehiclecruiseconfiguration.An entirewing with the Liebeck

LIOO3Mairfoil sectionat the designconditionshoulddevelopan estimated

three-dimensionalift coefficient(CL) = 1.6. With a 540 squarefoot wing

area and an approximate750 pound total vehicleweightat the start of cruise,

the vehicle'sair speed would be about 90 ft/secor 53 knots. Based on a mean

aerodynamicchord of 5.24 feet,Reynoldsnumberis 4.0 X lO5.

AircraftParasiteDrag Reduction

Parasitedrag reductioneffortswere concentratedon the wing and the

520 feet of strut guy wire runningperpendicularlyto the free streamvelocity.

Since cruiseperformancegreatlyaffectedvehiclecost effectiveness,all

analysiswas done assumingcruiseconfigurationat 60,000feet.

Reference3 calculatedHAAP vehiclewing parasitedrag coefficient

(CDp)= O.Oll3. The wing evaluatedconsistedof a high camberMiley airfoil

(reference7) for the i,boardwing and a sym!_etricsection for the outboard

foldingspan. The Mil.,yairfoilis much like the WortmannFX74-CLSo140alreudy

discussed. The Liebeckairfoilwith CDp = .009 has 20.4 percent lesswing

A-8

{> ,_,
parasitedrag. At designcrulse conditions,this resultsin a drag decrease

from 5.5 Ibf to 4.38 Ibf.

Parasitedrag reductionfor the HAAP vehicle'sstructuralsupportguy

wires has great potentiallargelydue to the high drag associatedwith the

wire's circularcylindershape. With a 0.036 inchdiameter,the wire has a

cruiseReynoldsnumberof 200. Reference3 reportsCD for a cy_,indricalwire

= 1.25,due largelyto pressuredifferencesbetweenfront and rear surfaces.

Adding 20 percent for wing-guywire interferenceeflects resultsinCD = 1.5.

With this relativelyhigh drag coefficientfor the 520 feet of wire, a drag

force of 2.11 Ibf results. .=treamlinedwire similarto that used in bi-

: plane applicationsoffers substantialdrag reduction. A streamlinedwire

with a finenessratio(:hordto maximum thickness)of 3.0 has CD = .25 (ref.8).

Typicalcrosssectionprofile.=are shown in the reference. A streamlined

_T_ wire will also have a smallerfrontal thicknessthan the diameterof the

originalcylindricalwire. In addition,it may be possibleto reducethe

-_ numberof wing supportingwires, due to the largemaximum thicknessof the

Liebeckairfoilwhich allows for the use of a torsionallystifferwing spar.

_, The preliminaryHAAP designconcept includedtwo wing supportwires above

k_ and two below each span. The high numberof v,ireswas originallyfelt neces-

_ sary to controlthe torsionalstiffnessof the wing. With the strongerwing

• spar, four out of eight of these 26-footwires could be eliminated,about a

_ 20 percent reductionin total wire length.

_ The resultof these designimprovementsis a substantialparasitedrag

; reduction. With a drag coefficientof .30 (.25plus 20C_Cfor interference

"t@ ?, effects),a streamlinedwire frontalthicknessof 0.03 inches (reduced]6.67

percentfrom 0.036 inch circularcylinderdiameter),and the 20 percent

_- reductioni_ totalwire length,the resultingwire parasitedrag = 0.279 Ibf

_. (86.8'/reduction). . Together,the wing and strut drag reductionanalysislowered

_ the vehicledrag 2.95 Ibf, from 7.61 Ibf to 4.66 Ibf. This increasedaerodynamicefficiencyrepresentsa 25.5 percent reductionin the total vehicle

parasitedrag, and lowersthe airplane'szero lift drag coefficient(%0)

from 0.0238 (ref. 3) to 0.01773.

i_. A-9

F
t'

Induced Dra_ and Wing OPtimization Anal_ysis

With the higher lift capability of the Liebeck LIOO3K.airfoil, the

addition drag induced by the wing at high CL'S must be considered in determining the best lift coefficient in cruise. Simply put, are the benefits of

higher lift outweighed by the large increase in induced drag (proportional to

the square of lift) that results? For best vehicle range, the L/D ratio is

i, optimized. Figure 6, a plot of L/D vs• CL, shows a CL of l.O to l• l to be

i optimal for best range performance with a maximized L/D ratio. L/D falls off

i rapidly at higher CL'S, showing the effects of greater induced drag. The HAAP

; vehicle, however, sees improved cost effectiveness with endurance or maximum

time in cruise optimization. Range is unimpoFtant in its mission since it

_:. will be flying a stationary pattern over its relay area. A plot of the

;, endurance ratio, CL3/2/C D, vs. CL in figure 7 shows a maximum endurance ratio

:_ of about 34 asymptotically approached at maximum CL. This effectively says that

i_ for maximum endurance and minimum power, the wing should be flown at its highest

°ii_" :_- design lift coefficient. For tile airfoil selected, the maximum threem

dimensional CL possible is about 1.6.

'_ CL

_- However, to fly the HAAP vehicle at a cruise of 1.6 requires the entire

wing to be of the high performance Liebeck wina section. This presents a potential

i._ problem in the climb configuration, since the outboard wing i_ folded over and

_ flying in an inverted position until deployment at 40,000 feet Flying at the

'i_ same angle of attack as the inboard section, but in the upside down position, the

:_ section would produce a high negative lift that might make the 40,000 foot deploy-

_ ment altitude unreachable. A solution exists in turning the outboard wing section

i'_ at the inboard/out'.)oardhinge point as the section is folded over and inverted.

}T This is mechanized by appropriately choosing the hinge line This turning or twist

_" at the hinge li:lewould result in the inverted section seeing lower angles of

_. attack, resulting in lower neg,_tivelift forces. The inboard wing could still

i_i_ fly at its design angle of attack an(_high CL with a less deleterious effect on !:

'_' climb perforr,lancefrom the inverted outboard section. For best climb performance,

negative lift effects need to be minimized; however, reference 6 shows dramatic

increases in airfoil parasite drag below CL'S = 55. The optimum hinge line angle

w,.)uldresult in a low enough negative lift coefficient without large parasite drag increased. Figure _ d(;ten:linesthis optimum turn angle to be

about lO degrees of twist where the endurance ratio is maximized at II.66.

A-IO

..... .................... . _ , • .._ ,-._-_._:-.j.................... . . ,._-
A largerturn anglewould resul_in great increasesin parasitedrag while a

smallerangle would allow too high a negativelift (:oeffia_epi.The . inverted

outboardwing sectionwould fly at an effectiveCL _ -.55 witl;a resulting

totalwing effectiveCL between .65 and .725and L/D = 13.7. These values

comparefavorablywith the climb performanceof the cost sensitivityanalysis

referencevehicle. It climbedat CL = 0.8 and L/D = 14.

Effectof Aero EfficiencyImprovementson Cost Effectiveness

To determinethe influenceof improvedaerodynamicperformanceon the

baselinereferencevehicle'scost effectiveness,a mission was flown changing

only the aero parameters. The aero-improvedreferencestandardincorporated

a climb CL = .65,climb L/D = I0.35,and a cruiseCL and 32 percenthigher

_- cruiseL/D. All other parametersremainedunchanged• Cost effectivenessfor

thismodel improved25 percent,from $.103 to $.077 per pound payload-houron

i_ station.

_ With the improvedaerodynamicefficiencyof the vehicle,greaterweight-

!_ carryingpotentialis realized. A secondaero-improvedmodel was flownwith

_ 200 poundsadditionalweightand an increasedmean aerodynamicwing chord of

:_ 7.0 feet. The increasedchord createdmore wing area, keepingwing loading

:_ nearlyconstant,and helpedmaintainReynoldsnumberflow closer to the airfoil

design condition. The lO0 poundseach in additionalfuel and payloadresulted

in a furtherimprovementof the cost effectivenessto $ 041 oer pound payload-

:_ hour,a 60 percent improvementfrom the $.103 value of the referencevehicle.

i--

i-=

_-11
4!r" "t)

j_

RECOMHENDATIONS

I Through the course of this analysis, many areas of additional needed

research and study were identified. A larger vehicle may prove to be more

cost-effective. Higher Reynolds number flow for better airfoil performance,

increased engine efficiency, and lower operating costs per payload poundhour are all advantages of a larger system that should be traded against

higher weight and cost. In addition, much work can be done to improve the

efficiency of propellers for this application. 1,

!'

!.

I

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I'

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?,.

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4"

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A-12

!,
REFERENCES

I I. Akkerman, d. W. and Taub, W. 14. Data Package for Airborne Experiment

1 Program Vehicle, Johnson Space Center, December "1979through February 1980.

2. Taub, W. H. and Akkerman J. W. HAAP-2 (High Altitude Airplane Platform),

I 3. Joh Turriziani nsonSpaceR.V. CenterEvaluation Drawing,January of an Airborne 1980. ExperimentProgramVehicle

Concept and Designed by the Johnson Space Center (W.O. J327), Hampton

Technical Center Memorandum V-IglOO/OLTR-132, May 23, 1980.

4. Perez, L. F. HAAP-2 Aerodynamic Evaluation, Entry Aerodynamic Design

Section, Johnson Space Center, June 2, 1980.

5. Wortmann, F. X. "The Quest for High Lift," AIAA Paper 74-1018, September

1974.

6. Liebeck, R. H. "On the Design of Subsonic Airfoils for High Lift,"

, AIAAPaper76-406,July1976.

!!i: 7. Miley, Reynolds S.Numbers," d. "An Analysis Mississippiof theStateDesign College, of Airfoil January Sections 1972. for Low

,i_, 8. Mises, R. V. Theory of Flight, Dover Publications, Inc., New York, 1959.

iS"

Jj__

'_i- o '

if

! .

t

._, i'

&-13

A
BIBLIOGRAPHY

Ayers, T. C. "SupercriticalAerodynamicsWorthwhileOver a Range of Speeds,"

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McCormick,B. W. Aerodynamics_Aeronautics_and FlightMechanics.New York;

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:!_ McMasters,J. H., and Palmer,C. M. "PossibleApplicationsof SoaringTechi_" nologyto Drag Reductionin PoweredGeneralAviationAircraft,"NASACR-

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-_ Miley, S. J "On the Designof Airfoilsfor LOW ReynoldsNumbers,"AIAA

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_-._ Patrick,T. J. "AerofoilsDown to CriticalReynoldsNumbersand the Per'- .

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Phillips,W. H. "Some DesignConsiderationsfor Solar-PoweredAircraft,"

: NASA TechnicalPaper 1675,June 1980.

Pick, G. S., and Lien, D. A. "TheDevelopmentof a two-Dimensional,High

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ii " ,, ........ ,r
i

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Whitcomb,R. T. "Reviewof NASA SupercriticalAirfoils,"ICASPaper 74-I0,

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Wortmann,F. X. "A CriticalReviewof the PhysicalAspects of AirfoilDesign

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ORIGI?IAL P_l.Ggli_

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THE HAAPOPERAT]NGPROCEDURE

:,_ TTA-T-2PS58 ]/6/83

i_ DOC, NO. DATE .,

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'..i_: eef.e,le, mY: James H. Akkerman(L_.-_ _ #/,)/1,_-

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i- APPROVEmyD: n.'auu_. Sve_lkovsk.y _ ':'.:, r,;'.". '_l,I.._ _._

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THERNOCHEMICATEST L AREA oF 16

1.0 Verify connection of system per the attached schematics (12 each):

i I, Test article air/exhaust

2, Test article oil

3. Test article fuel

i 4. Dynamometer

5. Cooling air control

ii 6. Throttle/choke 7. Turbocharger backpressure control

! 8. Facility vacuum/inbleed/dehumidifier

i 9. Facility safety

" 10. Coolingwater

II, Facillty cooling

!_', 12. Instrumentation/ignition/auto shutdown

2,0 Verify the following:

IF I. 0ii level in engine reservoirat top fin _I/2".

2. Oyno oil level within 5" of top of drum.

i;:?- 3. LN supply to facilitycooling shroud at 20" level or more•

:i_- 4. LE_'s calibratedfor alarm at 0.25 LEL (alarm funct.)

_:; 5. Both vacuum pumps operating bu_ valved "OFF"

6. Gasoline full to within 2" of top of sight glass used for

;_ volu_e measurement (I00 tO 130 octane) aviation fuel•

i_. 7. All instrumentationcalibratedand in date, zero, span and

functional.

,i'_ 8. Air control in "OPEN" position (schematic5)

::i g. Backpressurecontrol in "OPEN" position (schematicB).

r_ 10" Throttleclosed, choke open (schematic6).

_o II. Dyno control valves (#4 & 5) closed (schematic4).

_'_ 12. Air inbleed valves (#6 & 7) open (schematic8).

13. LN2 main shut off valve 733 closed (schematic]l).

i_. 14. LN? control valve (#10) dehumidifierc_osed (schematic8).

;i_ 15. Ba[tery chargedand charger "ON".

i'_ 16. Coolingwater "ON" and flowing5 gpm minimum (schematic11)•

" 17. Cooling fans connected and running (6 blowers, I fan;

,_;c schematic Ill

_ 18. Verify fan current 12+! amps.

Ig. Verify safety chain for chamberdour available (schematic9),

20. Verify C02 fire extinguishernearby.

'!_'> 21. Verify manifold pressure switch functionalat: 2 psig nlaxh,um•

_'" 3.0 Open fuel valves 12, 13, and 14 at bottom of measurement5igl}tglass

to allow fuel to till carburator (schematic3).

%, :: 4.J Open valve POVIUO4 to small vacuum pu:np(schumatic_) •

|l| Form 3|0¢ (Roy Mmr OO) PlAIw4-JSC 
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5.0 Open dyno load valve 4 (needle valve) about 2 turns (schematic 4).

6.0 Open throtle about _ turns, pull choke momentarily,and push "START"

button. Choke as necessary until smooth runntng is achieved.

Verify oil pressure and set rpm at about 2,000 with about 400 ps_ on

load gage (schematic 4).

7.0 Close chamberdoor and place _afety chain across the face of ti_e

door.

' 8.0 Close inbleedvalve HV7 to chamber (the one to the cold box). The

i altitude should 9o to about 5,000 ft (I0 psla) (schematicB).

9.0 Open the LN_ supply valve 733 and the valve to the cold box IIVIO

(schematic8).

(i_ I0.0 Reverify instrumentationready, all safety items checked,chain in _,_ place, LEt's not over 5%, and fire extinguisherhandy.

!!:

I!L II.0 Begin to add throttle, but control speed with load until wide-open

i,_'7 throttle is achieved at 2500 rpm. This should be below the speed

il,_,- where any boost Is achieved.

,_ 12.0 Close the air inbleedvalve HV6 (ambient)to let about B psia in the

!i.-'._ chamber, reducingload to allow speed to increaseto _rovide about

one atmosphere (in the engine air receiver-manifoldpressure).

i,_ 13.0 As the chamber temperaturedrops, begin to close off the ambient air

_!7. inbleed,HV6, and open the dehumidifierinbleed,HV? to supply dry

° air.

14.0 Continue to close off air inbleeduntil about 5 psia where the large

pump must be opened valve POV 1005 to the chamber,shnultaneously

•_'_ openingthe dry alr inbleed (HV7) to hold 5 psia. _hi~

_!_ 15.0 Continue to restrictthe Inbleed,holding the manifold pressure by

reducingthe load and allowing engine rpm to increase (maximum

_ engine rpm is 8,000).

"N'

'_ CAUTION: --Go slowly and watch cylinder heat temperature--itshould

_.,, not go over 425°F.

_!',o --Also, oii temperaturenot over 375"F.

.... L-_ Control L.N,flow to ke_p the cold wall at about -IU()%$

',i_ --Controlth_ dehumidifierLH_ to not waste LH? out the

,' dump line. "

i!7 --It LEt gets to 0.25, stop and tlx LduSe. il

,>

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NOTE: If cylinder head temperature goes over' 425"F, reduce manifold

pressure by slowing engine rpm, or by going to lower chamber

pressure without Increasing engine rpm.

16.0 At some cl_amberpressure, hopefully 1 psia, the manifold pressure

will reduce and the engine will slop. At this point, begin the

inbleed (valve HV7), shut off the large pump, and alluw tile cl_amber

to go to about 10 psia. Restart the engine and repeat as necessary

to define the engine operating limits.

NOTE: The air centre! may have to be used to regulate cylinder head

temperature tn the range between 8 psia and 4 psia

(schematic 5.)

:_ The exhaust backpressure control will have to be used to

balance the inlet air pressure (chamber pressure) and the

exhaust pressure to equal levels (schematic 7).

'_ 17. Repeat runs as required to get data at various pressure levels

(maximumand minimum power, SFC, etc.) i.e., 2, i.5, 1, 0.75 psia,

if possible. Keep the chamber as cold as possible without freezing

the Instrumentationlines.

18. After completion of test runs, shut off the LN_ (valves 733 and

_. HVlO), open the inbleed valves HV6 and HV7 and'allow the engine to

_ idle at about 1800 rpm and 600 psi load until the temperatureii_the

chamber returnsto about /O'F. .

J_ 19. Shut off the fuel (valves#12, 13, and 14) and allow the engine to

:_ deplete the supply in the plumbing before turningoff the ignition.

_ Let the instrumentationoperate for a cycle or two after engine

i_ shutdown.

i_ 20. Secure the system as follows:

it:

_ I. Turn off the coolingwater.

i_ 2. Turn off GN2 supply. See facilityschematic.

_ 3. Turn off vacuum pumps.

4. Turn off fans.

_ 5. Disconnectchain door restraint.

6. Power down the instrumentation.

7. Close load valves HV4 and HVS.

"_ B. Turn off battery charger.

J 21. Place catch pa, under door joint and open dour.

i ....... ,i i

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EMERGI!NCPRY OCtiOURL

A. In ca_¢+of fire in chamber during run:

1, Close inbleed valves HV6 and HV/.

2. Shut off fuel (HVt2, 13, and 14).

3, Monitor temperature,

B. In case of fire in facility area:

1, Call fire department x3333.

2, Use CO_ fire extinguishers.

3, Terminate facility power.

4. Close all doors,

5. Avoid smoke inhalation.

22, For each test run (to altitude and back) designate a run number, d=te,

etc.

23. Recheck data system calibration, etc. Record final "0" valves witl_

the system "down" before instrumentation is shutdown.

24. Record on data sheet any observations which may need work before the

next run.

25. In ca_e a measurement of specific fuel consumption is to be made,

close valve HV12 and open valve t4V13allowing engine to operate using

fuel out of sight 91ass. Measure and record on Fluke print the time

required to use the marked amount, reverse valves to resupply the

sight glass.

lie @om|ilOC(Royliar Illi) M_.JS_

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