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.

INVESTIGATION O F

THE

STATIC STABILITY CHARACTERETICS

OF

FIVE HWERSONC MISSILE CONFIGURATIONS AT

MACH NUMBERS FROM 2.29 TO4.85

By

Kenneth

L.

Turner

and

W . E. Appich, Jr,

Langley Aeronautical Laboratory

Langley Field, Va.

-

--

2

(NASA

Cii OR

TA4X OR AD NUMBER

(CATEGORY)

HI

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NACA RM

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NATIONAL ADVISORY

COMMITTEE

FOR AERONAUTICS

RESEARCH

MEMORANDUM

INVESTIGATION OF THE STATIC STABILITY CHARACTEBISTICS

OF FIVE KYPEBSONIC MISSIW CONFIGURATIONS AT

MACH

NIIMBmS

FROM 2.29

TO 4.65

By Kenneth

L.

Turner and

W.

H.

Appich,

Jr.

SUMMARY

An i nves t i ga t i on

wind t unne l t o determ

has been conducted

ne th e s t a t i c s t a b

i n th e Langley

l i t y c ha ra ct er

Unitary Plan

s t i e s of f i v e

hypersonic m is si le con fig ura tion s. The models te s te d were a basic body

with length-diameter ra t io of

10

and an ogival nose with

a

f ineness

r a t i o of

5 ,

the body with

a

10 f l a red a f t e rbody ( s k i r t ) , and the body

with two se ts of low -aspect-ratio cruciform fi n s.

An

ad di t io na l model,

known as the hypersonic t e s t ve hic le,

w a s

included t o s imulate

a

Langley Pi lot less Aircraf t Research Divis ion f ree-f l ight t e s t vehicle .

Te st s were performed a t Mach numbers of 2.29, 2.75, 3.22, 3.71,

and 4.65 and a t Reynolds numbers, based on th e body leng th , from

approximately 2.5 X 10 t o 15 x 10 .

6

The results show that there

i s

l i t t l e ef fe c t of Mach number, w ithin

th e t e s t Mach number range, on th e sl op e of t he norm al-force curve a t

low angles of a t t ack fo r the conf igura t ions t es t ed .

A

s k i r t of t he t ype

t e s t ed i s e f f ec t i ve i n p roducing

l i f t

and pitc hin g moment i n the t e s t

an gle -of -att ac k range. The use of

a

s k i r t , however, lead s t o a drag

penal ty wi th a corresponding lo ss i n l i f t - d r ag r a t i o . W ith t he cen t e r

of gravi ty

a t 50

percent of th e body length, the s ki r t ed and f inne d

models a re d i rec t iona l ly s t ab le a t the low angles of at tack.

A t

t he

hi gh er t e s t Mach numbers and a t the higher angles of a t tack, the di rec-

t i on a l s t a b i l i ty f or the f inned models becomes grea te r than th a t exper i -

enced a t angles of a t tac k near

0 .

at ta ck and low Mach numbers, th e fin ne d models tend toward i n s t a b i l i t y .

However, a t the high angles of

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2

INTRODUCTION

NACA

RM ~ 5 8 ~ 0 4

The design of hypersonic missiles i s , t o a larg e degree, d icta ted

by consi der atio ns of aerodynamic hea ting . Configuration s which have

sur faces tha t p resen t

small

angles t o th e ai rs t ream (e .g . , h ighly swept

l i f t i n g sur fac es) have been shown t o have comparatively low heating r at e s ,

and ar e th ere fore being considered fo r use as hypersonic ai r - to-ai r and

ground- to-ai r mi ss i le s . In order t o obtain more informat ion on such

co nfig ura tion s, th e Natio nal Advisory Committee f o r Aeronautics has

re ce nt ly undertaken an inv est iga tio n t o determine th e aerodynamic char-

a c t e r i s t i c s o f

a

family of mis s i le conf igurat ions . This inv es t i gat ion

i s t o be performed a t supersonic and hypersonic speeds, and

i s

t o cover

a la rg e Reynolds number range.

The models t o be in ve sti ga te d include

a basic body with length-

diameter r a t i o of 10 and an ogival nose with a f i neness r a t i o o f

5,

t he

body with

a 10

f l ar ed af terbody, and the body with two dif fe re nt se ts

of low-aspect-ratio cruciform f in s .

An

a dd it io n a l model, known as t h e

hypersonic t e s t vehicle ,

i s


a

Langley Pi lot less

Ai rcra f t Research Div is ion f r ee - f l ig h t t e s t veh ic le .

p rev iously t es ted i n the Langley 4- by 4-foot supersonic pressure tun-

n e l

a t

a Mach number of 2.01 and the r e s u l t s a re present ed i n ref ere nc e 1.

These models were

The present paper con t a i n s t he r e s u l t s

of

t e s t s made i n the Langley

Unitary Plan wind tun nel t o determine drag and s t a t i c long i tudin al and

l a t e r a l s t a b i l i t y c h a r a c t e r i s t i c s ob ta in ed

a t

Mach numbers of 2.29, 2.75,

3.22,

3.71,

and 4.65 and

a t

Reynolds numbers, based on the body length,

from approximately 2.5

X

lo6

t o

15

X

l o6.

Also

inc luded i n th i s paper

ar e comparisons of the da ta of t h i s report with data of reference 1.

SYMBOLS

The co ef fi ci en ts of f or ce s and moments ar e re fe rr ed

t o

the body

axes system.

A l l

aerodynamic moments a r e tak en abo ut t h e ce nt er of

gr av ity which

i s

located

a t

th e 50-percent length of the

missi le

being

t e s t e d .

Symbols used i n th i s paper a re as follows:

cA

Axial f o r ce

axia l - force coef f i c ien t ,

qs

Base ax ia l fo rce

base ax ia l - force coef f i c ien t ,

A, B CIS

C

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2

INTRODUCTION

NACA

RM ~ 5 8 ~ 0 4

The design of hypersonic missiles i s , t o a larg e degree, d icta ted

by consi der atio ns of aerodynamic hea ting . Configuration s which have

sur faces tha t p resen t

small

angles t o th e ai rs t ream (e .g . , h ighly swept

l i f t i n g sur fac es) have been shown t o have comparatively low heating r at e s ,

and ar e th ere fore being considered fo r use as hypersonic ai r - to-ai r and

ground- to-ai r mi ss i le s . In order t o obtain more informat ion on such

co nfig ura tion s, th e Natio nal Advisory Committee f o r Aeronautics has

re ce nt ly undertaken an inv est iga tio n t o determine th e aerodynamic char-

a c t e r i s t i c s o f

a

family of mis s i le conf igurat ions . This inv es t i gat ion

i s t o be performed a t supersonic and hypersonic speeds, and

i s

t o cover

a la rg e Reynolds number range.

The models t o be in ve sti ga te d include

a basic body with length-

diameter r a t i o of 10 and an ogival nose with a f i neness r a t i o o f

5,

t he

body with

a 10

f l ar ed af terbody, and the body with two dif fe re nt se ts

of low-aspect-ratio cruciform f in s . An a dd it io n a l model, known as t h e

hypersonic t e s t vehicle ,

i s


a

Langley Pi lot less

Ai rcra f t Research Div is ion f r ee - f l ig h t t e s t veh ic le .

p rev iously t es ted i n the Langley 4- by 4-foot supersonic pressure tun-

n e l

a t

a Mach number of 2.01 and the r e s u l t s a re present ed i n ref ere nc e 1.

These models were

The present paper con t a i n s t he r e s u l t s

of

t e s t s made i n the Langley

Unitary Plan wind tun nel t o determine drag and s t a t i c long i tudin al and

l a t e r a l s t a b i l i t y c h a r a c t e r i s t i c s ob ta in ed

a t

Mach numbers of 2.29, 2.75,

3.22,

3.71,

and 4.65 and

a t

Reynolds numbers, based on the body length,

from approximately 2.5

X

lo6

t o

15

X

l o6.

Also

inc luded i n th i s paper

ar e comparisons of the da ta of t h i s report with data of reference 1.

SYMBOLS

The co ef fi ci en ts of f or ce s and moments ar e re fe rr ed

t o

the body

axes system.

A l l

aerodynamic moments a r e tak en abo ut t h e ce nt er of

gr av ity which

i s

located

a t

th e 50-percent length of the

missi le

being

t e s t e d .

Symbols used i n th i s paper a re as follows:

cA

Axial f o r ce

axia l - force coef f i c ien t ,

qs

Base ax ia l fo rce

base ax ia l - force coef f i c ien t ,

A, B CIS

C

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NACA RM

L58DO4

3

l

cm

C

C

mO

Cn

cN

cN

U

cyP

2

M

9

R

S

XJ

Y

U

L h

P

Rol l i ng moment

r ol l i ng- moment coef f i ci ent ,

ssz

Pi t chi ng moment

pi t chi ng- moment coef f i ci ent ,

9s1

acm

au

l ope of pi t chi ng- moment cur ve,

pi t chi ng- moment coef f i ci ent at zero normal f orce

yawi ng- moment coef f i ci ent ,

Yaw ng moment

9s

2

sl ope of yaw ng- moment curve, cn

aP

Nor mal f or ce

nor mal - f or ce coef f i ci ent ,

ss

ac

au

l ope of nor mal - f or ce cur ve, .

Si de f or ce

si de- f orce coef f i ci ent ,

qs

3CY

sl ope of si de- f or ce cur ve,

aP

m ssi l e l engt h, i n.

f r ee- st r eamMach number

f r ee- st r ean dynam c pr essur e, l b/ sq f t

Reynol ds number

maxi mum cr oss- sect i onal area of t he cyl i ndr i cal body, sq f t

coor di nat es of nose of m ssi l e measur ed f r ompoi nt unl ess

ot her w se noted), i n.

angl e of at t ack of m ssi l e cent er l i ne, deg

t unnel f l ow angl e, deg

angl e of si desl i p of m ssi l e cent er l i ne, deg

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NACA RM L58mJ-1

APPARATUS AND

METHODS

Tunnel

The test s were perf ormed i n t he hi gh Mach number t est

sect i on of

t he Langl ey Uni t ary Pl an w nd t unnel , whi ch i s a var i abl e pr ess me,

cont i nuous- f l ow t ype. The t est sect i on i s 4 f eet squar e and appr oxi -

mat el y 7 f eet l ong. The nozzl e l eadi ng t o t he t est sect i on i s of t he

asymmet r i c sl i di ng- bl ock t ype whi ch per m t s a cont i nuous var i at i on of

Mach number f r omapproxi matel y 2.29 t o 4.65.

Model s

A dr aw ng

show ng t he f i ve model s t est ed i s pr esent ed i n f i gur e 1

and tabl e I gi ves t he geomet r i c char act er i st i cs of t hese model s.

The m ssi l e conf i gur at i ons wer e

Model I - body al one l engt h- di amet er r at i o of 10)

Model I1 - body wi t h

loo

f l ared ski r t

Model I 11 - body wi t h cruci f or m 5 del ta f i ns

Model

IV

- body wi t h cruci f or m 15' del ta f i ns

Model

V

- hyper soni c t est vehi cl e

The f i r s t

t he poi nt

whi ch i s

f our model s i ncor por at e a cyl i ndr i cal body wi t h an ogi ve nose,

of whi ch has a 0. 3- i nch r adi us of cur vat ur e. The f i f t h model

somewhat l onger ) has t he same cyl i ndr i cal port i on of t he body

but i t has a modi f i ed ;on K nose, t he poi nt of whi ch has a

0. 05k- i nch r adi us of cur vat ur e. Thi s model al so i ncor por at es a 10 s k i r t .

Hencef ort h, t hese model s wi l l be r ef er r ed t o as model s I t o V. The

model s ar e of st eel const r uct i on except f o r t he nose por t i on of model V

and the f l ared ski r t s whi ch wer e made of an al um num al l oy.

of model I 11 as i nst al l ed i n t he t est sect i on i s pr esent ed as f i gur e 2.

A phot ogr aph

For ces and moment s wer e measur ed by means of an i nt er nal l y mount ed,

si x- component , st r ai n- gage bal ance.

Test Condi t i ons and Procedure

The t est s wer e per f ormed at Mach number s of 2.29, 2.75, 3.22, 3.71,

and 4.65.

Mach numbers except 4.65, at whi ch Mach number i t was al l owed t o r i se

t o

-20

F.

The dewpoi nt t emper at ure was mai ntai ned bel ow -30 F f or al l

The s t agnat i on t emperature was mai ntai ned at appr oxi matel y

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NACA RM L58DO4

5

140

F a t a l l t e s t Mach numbers except

4.63,

a t which Mach number

it

was

held

a t

approximately 175O

.F.

The

following

t ab l e p r e s en t s t he t e s t cond it ions

of

each model:

Model

I

I 1

I 11

IY

v

Nominal

angles of

a t t ack ,

deg

-2 t o 25

-2 t o 25

-2 t o 25

0 , 7, 14,

and

20

-2

t o

25

0 , 7, 14,

and 20

-2

to 25

Nominal

angles

of

s ides l ip ,

deg

0

0

0

-3 t o

12

0

-3 t o 12

0

Mach

number

2.29

2.75

3.22

3.71

4.65

2.29

2-75

3.22

3-71.

4.65

2.29

2.75

3.22

3-71

4.65

2.29

2.75

3.22

3-71

4.65

2.29

2.75

3.22

3-71

4.65

R

12.5

x

io6

12.5

12.5

12.5

12 . 5

6

5.0 and 12.5 x 10

5.0 and 12.5

5.0

and

12.5

5.0

and

12.5

7.5

and

12.5

2.5, 5.0, and 12.5

X

106

2.5, 5.0, 12.5

2.5, 5.0,

and

12.5

2.5, 5.0,

and

12.5

5.0,

7-57

and

12.5

6

12.5

x i o

12.5

12.5

12.5

12.5

15.0 x i o

15.0

15.0

6

15.0

15.0

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6

CN . . .

CA

. . .

cm

. . .

c2

. .

n

. . .

cy

. . .

a, deg

.

P? deg 0

M . . . .

CORRECTIONS

AND ACCURACY

Accuracy a t -

R = 5 X l O

66

R

=

12.5

x

i o

or 15 x 106

6

R

= 2.5 x

i o

a.

34

+o .067

20.029

20.007 fO 003 +o .002

k0 .055 +o

.027

a 011

k0.004

fO 002 XI.

001

k0 .055 +o

.027

20.011

+O

.134

20.067

+o .029

20. LOO

+o

. l oo

+o

.loo

kO.100

+o . loo +o . l oo

20.015 +O .015 +O .015

NACA RM L58Do4

The angles of a tta ck and si de sl ip have been corrected f o r defle c-

t i o n of th e balance and st in g under load.

In o rder t o ob ta in re l iable values of base ax ia l force , a base

block

w a s

f i t t e d secure ly t o the model

sting

wit h about 1/8-inch gap

between t h e block and th e base of th e model.

model was cy li nd ri ca l and of

t h e

same diameter as the base of the model

being invest igate d. (See f i g .

2 . )

Measurements were taken of the

pres sure ex is ti ng between th e base block and th e model base and thes e

pressure measurements were converted in to base a x i a l forc e. The

axial-

fo rce da ta

on t he

plots of aerodynamic coefficients are 'not adjusted

fo r ba se a x i a l f o r c e . In order t o ad jus t these da ta, the base ax ia l

fo rc e fo r a given model a t a g iv e n a t t i t u d e

and Mach number

must

be

subtracte d from the ax ia l forc e f o r th e same model

a t

t h e s a n e a t t i t u d e

and Mach number. During

t e s t s of

model V, fa ul ty equipment cur ta i led

base pres sure measurements, and base-axia l-force da ta f o r th i s model ar e

not presented.

The base block f o r each

The accuracy of t he i ndi vid ua l nieasured qu an ti ti es , based on ca li -

b ra t ion and repe a tab i l i ty of da ta , i s es timated t o be wi th in the f o l -

lowing

l i m i t s

:

-

._

- -

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NACA RM L58Do4

M

2.29

2-75

3.22

3.71

4.65

7

, deg

0.40

30

. 10

30

95

-

Cal i br at i on of t he t unnel t est sect i on has not been compl et ed.

Measur ed pr essure gr adi ent s ar e suf f i ci ent l y smal l , however , t o assur e

negl i gi bl e cor r ect i ons due t o model buoyancy ef f ect s.

The dat a have not been cor r ect ed f or f l ow angul ar i t y. These cor -

r ect i ons at t he cor r espondi ng Mach numbers are i ndependent of model

angl e of at t ack and ar e as f ol l ows:

PRESENTATION OF RESULTS

The r esul t s of t he i nvest i gat i on ar e pr esent ed i n t he f ol l ow ng

f i gur es:

Fi gur e

Typi cal schl i er en phot ogr aphs of model I1

. . . . . . .

. .

. . 3

Typi cal schl i er en phot ogr aphs of model

I11

.

.

. . . . . . . . . 4

Var i at i on of base axi al - f or ce coef f i ci ent wi t h angl e of

attack

. . . . . . . . . . . . . . . . . . . . . . . . . . . .

5

Ef f ect of base bl ock on aer odynam c char act er i st i cs

of

model 111

. .

.

.

. . . . . . .

. .

.

.

. .

.

. . . . . . . . 6

Aer odynam c char act er i st i cs of model I n pi t ch. j3

=

Oo

.

.

. .

7

Aer odynam c char act er i st i cs of model

I1

i n pi tch. j3 =

0

. . . . 8


I11

i n pi tch. j3

=

0 . . . 9

Aer odynam c char act er i st i cs of model I V i n pi tch. 10

Aer odynam c char act er i st i cs of model V i n pi t ch. j3 = 0

. . . . 11


I11

i n si desl i p.

a

=

0

.

.

12


I11

i n si desl i p.

Aer odynam c char act er i st i cs of model I11 i n si desl i p.


Aer odynam c char act er i st i cs of model I V i n s i des l i p . . . . . .

16

j3

=

o

.

. .

.

~ = 7 . 2 O. . . . . . . . . . . . . . . . . . . . . . . . . . . 13

a = 1 4 . 6 . . . . . . . , . . . . . . . . . . . . . . . . . . . 14

~ = 2 0 . 9

. . . . . . . . . . . . . . . . . . . . . . . . . . .

15

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8

Figure

Summary of longi tu dinal s t a b i l i t y ch ar ac ter is t ic s of the

f i ve m is s i l e configura t ions

. . .

.

.

. . . .

.

. . . .

. .

17

Summary

of l a t e r a l s t a b i l i t y cha r ac t e r i s t i c s of model

I11 . .

18

Summary of l a t e r a l s t a b i l i t y cha r ac t e r i s t i c s of model I V

. . 19

DISCUSSION OF FBSULTS

Effect of Base Block

In order t o de termine the e f fe c t of the base b lock-on the

s t a b i l i t y

ch ara cte r is t ic s presented, model I11

was

te st ed with and without th e

base block. (See f i g . 6 . )

base block produces

a

pos i t i ve

i s

not mater ia l ly a l t e red .

t e s t r e s u l t s p re se nt ed .

The re su l t s show th a t the add i t ion of th e

s h i f t ,

bu t t he deg ree of s t a b i l i t y

The base blocks

were

i n p la ce

fo r

a l l other

C

mo

Longi tud ina l S tab i l i ty

The l ong i t ud ina l s t ab i l i t y cha r ac t e r i s t i c s o f t he f i ve mi s s il e s

ar e pre sente d i n summary form pl o tt ed ag ain st Mach number i n fi gu re

17.

It may be seen t h a t the normal-force-curve slo pes of a l l f i ve models a re

invar ian t w i t h Mach number

a t

the low angles of at tack.

A t

the h igh

ang les of at ta ck , however, th e normal-force-curve slope s decrease with

an in cr ea se i n Mach number.

It

may be noted t h a t t he increment of

normal-force-curve slope provided by th e f i n s and s k ir t s

i s

e s s e n t i a l l y

in va ria nt with Mach number and th a t t he decrease i n normal-force-curve

s lope noted a t high angles i s due t o l oss of l i f t on the body and not

on the f ins

o r s k i r t s .

O f the models tested

a t

th e low angles of at ta ck, th e f in ned models

(models I11 and IV) have th e g re at e st normal-force-curve slope, and th e

model without f ins or skir t (model I ) has the least normal-force-curve

slope. Model

I11

develops more

l i f t

th an model

IV,

as

would be expected,

from consideration of the geometry of the two models.

A comparison of model I1 and model I V shows that the

s k i r t

f o r

model

I1 i s

approximately as long a s th e f i n s of model

I V,

but the

leading-edge angle of th e fi n s of model

IV

i s much la rg e r. The da ta

indicate that model

I1

develops

s l i g h t l y l e s s l i f t and has more drag

than model

N.

S i mil a r r e s u l t s i n r e f e rence

2

point out that an increase

i n the leading-edge angle or length of

a

s k i r t

w i l l

increase the l i f t

developed by th e s k i r t . The use of

a

sk ir t , however, leads t o

a

drag

penal ty with a corresponding

10

-

ss i n l i f t - d r a g r a t i o .

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NACA RM L58D04

9

Figure 1.7 al so shows th a t some sta bi l i z in g device

i s

necessary

f o r model I a t the low angles

of a t t ack i n or der t o ob t a i n a longi tudi-

na l ly s ta bl e mis s i le i n the t e s t Mach number range wi th the cente r of

g r a v i t y a t

50

per cen t of th e body leng th.

Mach numbers s l i g h tl y below th e t e s t range, th e s k i r t used on modelV

may not be l arg e enough t o produce pos i t i ve longi t udina l s ta b i l i t y .

The lopg itud ina l s t a b i l i t y of t h e s ki r t ed models in crea ses somewhat

wi th in cr ea se i n Mach number a t the low and high angles of at tack,

whereas the long i tudin al s ta b i l i t y of model N decreases with increase

i n Mach number.

These da ta ind ica te tha t

a t

Model I11 has the l ar ge s t s t a t i c margin of th e models tes t ed .

A t

t he low angl es of a t t ack , t he l ong i t ud ina l s t ab i l i t y ch a r ac t e r i s t i c s f o r

model I11 are re la t i ve ly cons tant wi th var i at ion i n Mach number i n the

t e s t Mach number ra nge.

It may a ls o be seen i n f ig ur e 17 tha t the da ta ob ta ined a t t h e

(Ref. 1 dat a are indica ted by symbols i n

t e s t Mach numbers agree very we ll with da ta shown i n refe ren ce

1

a t a Mach number of 2.01.

f i g . 17.)

D i re c ti on a l s t a b i l i t y

Models I11 and

I V,

th e fin ned models, were th e only models te s te d

i n s i d e sl i p .

show the mi ss i le s , i n general , t o be di re ct ion al ly s ta ble . These f i g -

ures also show

a

no nl in ea rit y i n th e yawing-moment curve a t low side-

s l i p angles . This nonl i near i ty inc reases wi th angle of a t ta ck and,

i n some ins tances , the mis s i l es a r e d i re c t io na l ly unstab le fo r a small

s id es l ip r ange near 0 .

higher s ide s l i p angles . These f igu res also show l i t t l e change i n

l o n g i t u d i n a l s t a b i l i t y o r normal-force co eff ici en t throughout the angle-

of - s ides l ip r ange .

s ent ed i n f i gu r e s

18

and

19

f o r bo th f i nned mi s s i l e s a r e f o r s lopes

between +2O of si d es li p , and because of th e aforementioned no nl in ea rit y

i n the yawing-moment curves do not prese nt th e complete s t a b i l i t y pic tu re

Data presen ted i n f igu res 12 t o 16 fo r the f inned miss i les ,

This i n s t ab i l i t y d i sappear s , however, a t the

The yawing-moment and si de -f or ce der iv at iv e s pr e-

It

may be seen i n fig ur es

18

and

19

t h a t

a t

the lower angles of

a t t ack , 0 and 7.0 f o r model

I11

and 0 for model N, t h e d i r e c t i o n a l

s t ab i l i t y of t he f inned mi s s i l e s

dec rease s wit h in cr ea se i n Mach number;

however, a t the higher angles of a t tac k (14.5' and 2O.9O) t he d i r ec -

t i o na l s ta b i l i t y increases wi th increa se i n Mach number. It

i s

a l s o

i n t e r e s t i n g t o n o te t h a t a t th e hig her t e s t Mach numbers and a t

an

angle of a t t ac k of 20.9' th e di rec t io na l s t a b i l i t y fo r model I11 (and

t o a l im ited ex tent, model

N

ecomes greater than that experienced

a t angles

of

at tack near 0 . It

i s

be l ieved t ha t t he r ea son f o r t h i s

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10

NACA

RM L58D04

phenomenon

i s

t h a t a t the high angles of a t ta ck there

i s

an increase in

dynamic pressure on the lower f i n which increas es t he effe ctiv en ess of

t h e l ower f i n a t

a

g r e a t e r r a t e t h an t h a t

a t

which the upper f i n

i s

los ing e f fec t iveness .

and

I la

The pitching-moment-curve and norma l-force-curve slo pe s

of the se hypersonic mis sil es a t a Mach number of 2.01 have been

C

NCL)

obtained from reference 1 and a r e p l o t t ed i n f i gu r e s 18 and 19. Since

th e models a re

a l l

symmetrical, it

i s

perm issible t o compare C and

C

a t zero angle of s ide s l i p w i th

at ta ck . These slope s, shown i n symbol form i n fi gu re s

18

and

19,

show

excellent agreement with the data reported on herein.

ma

a t zero angle of

CnP and cyP

a

The da ta on the bas ic p lo t s ind ica te th a t the d ihedra l e f fe c t

i s

es se nt ia l l y zero f o r th e f inne d models through t he t e s t Mach number

and ang le-of-a ttack range.

Figures

18

and

19

indicate negat ive values of

C

a t a l l a ngles

of a t tack.

A

comparison

of

these two

f igures

shows t h a t model

I11

has

the la rg er negat ive values of

bas is

of th e d iff er en ce i n model geometry.

. This would be expected on the

cyP

A l l models other than I11 and

N

are symmetrical models and would

t he r e f o r e have i de n t i c a l l ong it ud i na l and l a t e r a l s t ab i l i t y cha rac t er -

i s t i c s a round

0

angle of a t ta ck and s id es l ip .

Reynolds Number Effect

The s t a b i l i t y da ta f o r models

I1

and

I11

a t Reynolds numbers of

I t i s

eas i ly seen tha t the p i t ch and s ides l ip curves

6 6

6

6

2.3 x 10 5.0 X

10 7.5

X 10 and 12.5 X 10 are shown in f i gu res 8,

9, 12, and 14.

have th e same r e la ti v e shape, re ga rd le ss of Reynolds number i n th e t e s t

Reynolds number range.

however, f o r th e thr ee t e s t Reynolds numbers, dependent on a t ti t u d e

and

Mach number. I t i s bel ieved, moreover, th at th e data taken a t a

Reynolds number of

12.3 X

10

s ince the balance loads a t t h i s higher Reynolds number a re i n the range

t o ob ta in accura te da ta .

Reynolds numbers t o check th ose take n a t t h e hig her Reynolds

number

i s bel ieved t o be e nt i r el y due t o balance accuracy.

e n ti t l e d Corrections and Accuracy. )

There

i s

a gen eral intermixing of data points ,

6

accura te ly def ine the s t ab i l i ty curves ,

The in a bi l i ty of th e data taken

a t

the lower

(See sect ion

8/20/2019 19710066227

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NACA

RM

L58DO4

CONC

WSIONS

11

The re su l t s of an inve s t igat io n of f iv e hypersonic mis s i le configura-

t i o n s

a t

Mach numbers of 2.29, 2.75, 3.22,

3.71,

and

4.65

and a t Reynolds

numbers, b ased on the body le ngt h, from approxim ately

2.5 X 10

15

X

10

indicate the following conclusions:

6

t o

6

1.

With th e c enter of gr av ity a t 50 perce nt of th e body length,

the sk i r t ed and finned models a re d i rec t ion a l ly s t ab le

a t

the low angles

of a t tack. A t

the higher t e s t Mach numbers and a t th e h igher angles of

a t t a ck , t h e d i r e c t i o n a l s t a b i l i t y

f o r

th e finn ed models becomes g re at er

than that experienced

a t

angles of a t ta ck near Oo.

high ang les of a tt a ck and low Mach numbers, th e fi nn ed models te nd

toward ins tab i l i ty .

However, a t t h e

2.

A s k i r t of t he t ype t e s t ed

i s

ef fe c t iv e in producing

l i f t

and

pitching moment

i n th e t e s t angle-of-attack range. The use of

a

s k i r t ,

however,

leads t o a drag penal ty wi th a corresponding l oss i n l i f t - d ra g

r a t i o .

3.

The model with the 5 O f i n s has the l a rg es t s t a t i c marg in and

normal-force-curve slope of th e models te st ed .

4.

There

i s

l i t t l e e ff e ct of Mach number on th e slope of th e

normal-force curve a t low angles of at tack.

Langley Aeronautical Laboratory,

National Advisory Committee for Aeronautics,

Langley Field,

V a . ,

March

19, 1958.

1.

Robinson, Ross

B.:

Wind-Tunnel Investigation

a t

a

Mach Number of

2.01 of th e Aerodynamic Ch m a ct e ri st ic s i n Combined Angles of Atta ck

and Sid es li p of S ev era l Hypersonic M iss ile Configurations With

Various Canard Controls.

NACA

RM ~ 5 8 ~ 2 1 ,

958.

2.

Lavender, Robert E.: Normal Forc e, P it ch in g Moment, and Center

of

Pressure of Eighty Cone-Cylinder-Fruskum Bodies of Revolution a t

Mach Number 1.50. Rep.

6R3N3

Ord. Missile Labs., Redstone Arsenal

(Huntsvi l le , Ala.), Apr. 5, 1956.

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12

TABLF: I.

MODEL GEOME?ITIC CHARACTEKCSTICS

NACA

FU

L58DO4

lody :

Length, in. . . . . . . . . . .

Cross-sect ional area , sq i n . .

Fineness ra t io of nose . . . .

Length-diameter ratio . . . . .

Moment-center location, per-

cent length

. . . . . . . . .

Diameter, in. . . . . . . . . .

ik i r t :

Length, in. . . . . . . . . .

Base diameter, in .

. . . . . .

Leading-edge angle, deg

. . . .

ase mea, sq i n . . . . . . .

' ins

:

Area exposed,

2

f i n s ,

sq

i n .

.

Root chord, in. . . . . . . . .

Tip chord, i n . . . . . . . . .

Span exposed, in.

. . . . . .

Span to ta l , in .

. . . . . . .

Taper ra t io . . . . . . . . . .

Aspect ratio, exposed

. . . . .

Span diameter ratio . . . . . .

Leading-edge angle, deg . . . .

odel

I

io. OC

3 . 0 ~

7-07

5.0C

-0.oc

j0.

OC

ode ?

I1

30.oc

3 . 0 ~

7.0;

5 . 0 ~

LO.

oc

jO.O(

6.0:

5.1:

20.6t

LO

O(

odel

I11

30.00

3.00

7-07

5.00

10.00

50.00

34.36

19.12

0

3.20

6 . 2 ~

0

0.26e

2.07

5

Iode1

rv

50.00

3.00

7-07

5 00

LO. 00

50.00

9.55

5.97

0

3 . 2 ~

6 . 2 ~

0

1.07

2.07

15

ode1

v

35

*

11

3.00

7.07

5.00

11.70

50 00

4.67

4.64

16.91

10.00

8/20/2019 19710066227

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NACA RM L’j8Do4

7.00

8.00

9.00

10.00

11.00

12.00

13.00

14.00

13

1.073

1.176

1.262

1.335

1.394

1.441

1.474

1.493

Tangency points

t - d 30ad .

models I , 11, I11

and IV noses

6.63’

p=fF

.30 .300

5 13

l oo

25.25 rad;

Model

I1

I 6.00 .963

15.00 I 1.500

6.20

A.

.094 rad.

Model I11

-1.60

-

6

.

-----.AI koord ina te s

f o r 1

Model

IV

X-

1 1 7 . 5 5

Model V

1.698

1.947

3.693

3.945

4.938

5.076

6.444

7.944

9.444

10.994

12.453

13.944

15.444

.768

.918

1.059

1.188

1.296

1.389

1.461

Figure

1.

Miss i le conf igurat ions tes ted.

All

dimensions

i n

inches

unless otherwise s ta ted. )

t

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14

8/20/2019 19710066227

17/91

NACA

RM

L58DO4

a =

- 2 . 4O

a = O o

a

=

2 . 0 °

a = 6 . 2 O a =

20.9O

a )

M = 2.29. L-58

180

Figure

3 .

Typical sch lie ren photographs

of model 11.

B =

Oo

8/20/2019 19710066227

18/91

16

NACA

RM

L58DO4

a

O o

a = - 2 .40

a

= 2 . 0 °

a =

6 . 2 O

(b)

M = 2.75.

Figure 3 . - Continued.

a

= 20.9O

I,-38-181

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NACA RM

L58DO4

i

a = - 2 . 5 0

a

O o

a

= 2 . 0 0

a = 6 . 1 °

a = 20.7O

(c)

M

= 3.22. L-58-182


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18

NACA

RM L38DO4

a =

- 2 .4

a = 0

a = 2 .1

a

= 6 . 1

d ) M = 5-71.

Figure 3 . -

Continued.

a 20.7

L-58-183

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21/91

NACA RM

L58D04

0

8/20/2019 19710066227

22/91

20

NACA

RM L58D04

a=-2.4O a

= O o

a

=

2.0°

a = 12.4O a

= 20.6O

a) M

=

2.29. L-38-18?

Figure 4. Typica l sch l i e ren photographs of

model

111.

p =

0 .

8/20/2019 19710066227

23/91

NACA

EM L58DO4

21

a

= 2 . 4 O a

= O o

a

= 2.O0

a

=

12.3O

(b)

M

= 2.75.

Figure 4.- Continued.

a

20.50

1-58-

186

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24/91

22

a

= -2.40

NACA RM L58D04

a

=

12 .2O

a O o

a = 2 . 0 °

( e ) M =

3.71.

Figure

4.-

Concluded.

a 20.2O

L-58-187

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NACA RM

L58DO4

A B

C

Fi gur e 5.- Variat ion of base ax ia l - force coef f i c ien t w i th angle of

attack.

f3

= 0 .

3

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24

NACA RM ~ 5 8 ~ 0 4

C

m

.6

.4

.2 c

0

'.2

a )

M

= 4. 63; l = Oo.

Figure

6.-

Effect of base

block

on aerodynamic chasacteris t ics of

model 111.

R = 7.5 x

106.

8/20/2019 19710066227

27/91

NACA

RM L58D04

Y

d e g

(b)

M =

4.65;

CL

=

Oo.

Figure

6.

- Concluded.

C

n

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NACA RM L58D04

Fi gur e

7. -

Aer odynam c char act er i st i cs

of

model

I

in

pi t ch.

p

=

Oo;

R =

12.5 x

106-

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29/91

NACA

RM L58D04

0 4 8 12

16 20

24 28

a , d e g

Figure

8.-

[email protected]

c h a r a c t e r i s t i c s of

model I1

i n

p i t ch .

f3

=

Oo.

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30/91

NACA €34 L58DO4

4

0

3

(a) M = 2.29.

Fi gur e 9.- Aerodynamic characteris t ics of

model

111 in p i t c h .

f3 = 0’.

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31/91

NACA

RM

L58m4

a , d e g

( b ) M

=

2.73.

Figure

9.-

Continued.

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32/91

NACA

RM

L58D04

c ) M

= 3.22.

Figure

9.-

Continued.

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33/91

(d) M = 3.71.

8

Figure

9.-

Continued.

8/20/2019 19710066227

34/91

NACA RM ~581x14

e )

M

= 4.63.

Figure 9.- Concluded.

8/20/2019 19710066227

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NACA

RM L58Do4

n

.

.

4

Figure

10.-

Aerodynamic cha r ac t e r i s t i c s of m o d e l I V i n p i t c h .

p =

00;

R

= 12.5 x 106.

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Cm

N

.4

4

NACA RM L58D04

G

Figure

11.

Aerodynamic ch ar ac te ri s t ic s of modelV in pitc h.

p

=

Oo;

R = 13 x l o6.

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NACA RM L58Do4

.8

.4

Cn

0

'

.4

-3 -2 0

2

4 6 8 IO 12 14

9

d e g

(a)

M

= 2.29.

Figure 12.- Aerodynamic characteris t ics

of

model 111 i n s i d e s l i p .

a = oo.

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NACA

RM L58Do4

-4 -2 0 2 4 6 8 IO 12 14

8 , d e g

(a)

Concluded.

Figure

12.-

Continued.

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39/91

NACA

RM L58D04

CY

(b)

M = 2.75.


.8

.4

Cn

0

.4

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40/91

(b)

Concluded.

Figure 12.

Continued.

NACA

RM L78DO4

8/20/2019 19710066227

41/91

NACA

RM L38DO4

-4 -2 0 2

8 IO I 2 14

( c ) M =

3.22.

Figure

12.

Continued.

.4

(

0

.4

8/20/2019 19710066227

42/91

NACA RM

L38D04

( c ) Concluded.

Figure 12. - Continued.

8/20/2019 19710066227

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NACA

RM

L58D04

.8

.4

Cn

0

-

.4

4

(a)

M =

3.71.

Figure 12.-

Continued.

8/20/2019 19710066227

44/91

NACA

RM L38D04

(d) Concluded.

Figure 12.

Continued

8/20/2019 19710066227

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NACA

RM

L58D04

43

C Y

-4

-2 0 2 4 6

8

IO 12 14

8 9

d e g

( e )

M = 4.63.

.8

0

Figure

12.

Continued.

8/20/2019 19710066227

46/91

NACA

RM

L58D04

-4

-2 0 2 4 6 a

IO 12 14

B d e 4

( e )

Concl uded.

Figure 12.-

Concl uded.

NACA

RM L58D04

8/20/2019 19710066227

47/91

. 8

.4

c n

.o

.4

(a)

M =

2.29.

Figure 13 . - Aerodynamic

c h ma c t e r i s t i c s

of model I11 in sideslip.

a

= 7.20; R

= 12. 5x

106.

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NACA .RM

L58Do4

-4

-2

0

2 4 6 8 IO I2 14

8 9

d e g

a ) Concluded.


8/20/2019 19710066227

49/91

NACA RM

L58DO4

4 -2 0 2

4

6 8 IO 12 14

B

d e g

(b)

M =

2.75.

Figure

1 3 . -

Continued.

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50/91

NACA RM

L58DO4

-4

-2

0

2 4 6 8 IO

I2 14

B ,

d e g

(b)

Concluded.

Figure

13 . -

Continued.

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51/91

NACA RM L58DO4

-4

-2 0 2 4

6 8

10 12 14

9 d e g

c )

M = 3.22.

Figure 1 3 . - Continued.

8/20/2019 19710066227

52/91

NACA RM

L58DO4

-4 -2

0 2

( c)

Concluded.

Figure 13 .

-

Continued.

2

3

NACA RM

~ 3 8 ~ 0 4

8/20/2019 19710066227

53/91

-4 -2 0 2 4

6 8

IO 12

14

8 9 d e g

.8

,.4

(d) M

=

3.71.

Figure

13 . - Continued.

8/20/2019 19710066227

54/91

NACA RM L58Do4

(d) Concluded.

Figure

13 . -

Continued.

NACA RM ~ 3 8 ~ 0 4

8/20/2019 19710066227

55/91

-4

2

0

2

4 6 8

10 12

B , d e g

.

(e)

M

=

4. 65.

0

-

-4

4

Figure 13 . -

Continued.

8/20/2019 19710066227

56/91

54

NACA

RM L58DO4

-4 -2 0 2

4

6 a

IO

12 14

8 ,

d e g

( e )

Concl uded.

Fi gur e 1 3 . - Concl uded.

NACA RM L58DO4

8/20/2019 19710066227

57/91

CY

.8

.4

Cn

0

-4

-2 0

2

4 6 IO I2 14

8 , d e g

a)

M

=

2.29.

Fi gur e

14. -


CG = 146.

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58/91

NACA RM L58Do4

(a) Concl uded.

Fi gur e

14. -

Cont i nued.

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-4 -2 0 2 4

6 8

I O

I2 14

8 , d e g

.8

.4

Cn

0

.4

(b) M = 2.75.

Figure 1-4.-

Continued.

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60/91

NACA RM

~ 3 8 ~ 0 4

( b ) Concluded.

Figure

14.-

Continued.

NACA

RM ~38~04

8/20/2019 19710066227

61/91

CY

-4

-2

0 2

8 IO 12 i

( c ) M = 3.22.

Figure

14. -

Continued.

0

-.4

4

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62/91

NACA RM L58Do4

( c ) Concluded.


4

NACA F N L58DO4

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.8

.4

Cn

0

-

.4

-4

-2 0 2 4 6 8 IO 12 14

B ,

d e g

d) M =

3.71.

Figure 14.- Cont i nued.

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NACA RM L58DO4

4 -2

0

2

4 6 8

IO

12

14

8 ,

d e g

(

d)

Concl uded.

Figure

14. - Continued.

NACA

RM L38D04

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65/91

-4

-2 0 2

4 6 8 IO

12 14

8 , d e g

.8

.4

Cn

0

- .4

(e ) M = 4.65.


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64

NACA

RM

~ 7 8 ~ 0 4

4 -2 0

2

4 6 a to

I2

14

8 9

d e g

(e) Concluded.

Figure

14.

Concluded.

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Fi gur e

15.-

Aer odynam c char act er i st i cs

of model I11 i n

si desl i p.

a

=

20.90;

R

=

12.5 x

106.

.8

.4

Cn

0

.4

66

NACA RM

L58D04

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(a) Concluded.

Figure

15.

Continued.

NACA RM L38DO4

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CY

-4 -2 0 2 4 6

8

10

12

14

B

d e g

.8

.4

Cn

0

.4

(b)

M

=

2.75.

Figure 15 Continued.

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-4

-2 0 2

4

6 a 10

12

14

B ,

d e g

(b)

Concl uded.

Fi gur e

15.-

Cont i nued.

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-4

-2 0 2

4

6 8

IO

I2 14

P 9 d e g

( c ) M =

3.22.

.8

.4

Cn

0

4

Figure 15.- Conti nued.

NACA RM

~581x14

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4 -2 0 2 4 6 8 IO

I2

14

B d e g

c) Concl uded.

Fi gur e 15.- Cont i nued.

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CY

4

-2 0 2 4 6 8 IO 12 14

B

d e g

(d )

M

=

3.71.


.8

.4

Cn

0

.4

72

NACA RM ~ 3 8 ~ 0 4

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d ) Concluded.


NACA RM

L58DO4

7

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4 -2

0

2 4

6

8 I O I2 14

B , d e g

( e ) M = 4.65.

.8

.4

0

.4

Cn


74

NACA

RM

L58DO4

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-4 -2

0 2

4 6 a IO 12 14

B 1 d e g

(e) Concluded.

Figure 15.- Concluded.

NACA

FM

L58DO4 75

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2

Y

-4

-2

.4

0

.4

0

2

8 I O

I2

14

a )

M

=

2.29.

Figure 16.-

Aerodynamic characteristics of model

I V

in

s i d e s l i p .

R

= 12.3 x lo6.

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(a) Concluded.

Figure

16.- Continued.

3

f

NACA RM L58DOk >

77

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CY

8 9

d e g

(b) M = 2.75.

Figure

16. Continued.

.4

o

Cn

-

.4

4

78

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-4 -2 0 2 4 6 8 IO I2 14

8 ,

d e g

(b)

Concluded.

Figure

16.

Continued.

79

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4

-2

0

2 4 6

0 IO 12

14

8 9

d e g

c)

M =

3.22.

.4

o

Cn

.4

Figure

16.

- Continued.

80

NACA RM ~ 3 8 ~ 4

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-2 0 2 4 6 a I O 12 14

8 , d e g

( e )

Concluded.

Fi gure 16. - Continued.

NACA RM

~ 3 8 ~ 0 4

81

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-4

-2

0

2 4 6 8

IO

12 14

8 , dell

.4

0

.4

-l

(d)

M =

3.71.

Figure

16.

Continued.

82

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-4 -2 0 2 4 6 8 IO I2 14

B d e n

(a)

Concluded.

Figure

16.

Continued.

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CY

-4 -2

0

2 4

6

8 IO 12 I

8 , d e g

( e ) M = 4.65.

Figure

16. Continued.

.4

0

- .4

4

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e) Concl uded.

Figure

16.

Concl uded.

85

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.o

.o

C

ma

. o

.o

-

.o

M

( a ) u = O.

. 8

Figure 17.- L o ng it ud in al s t a b i l i t y c h a ra c t e r i s t i c s o f t h e f i v e mi s s i l e

configura t ions

.

86

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M

(b)

CY = 6O t o 20 .

Figure 17.

-

Concluded.

> 3

NACA RM L58DO4

87

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B

M

Figure 18.- L a t e r a l s t a b i l i t y c h a r a c te r i st i c s

of

model

111.

88

NACA

RM

L58DO4

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.

M

Fi gur e

19.

L a t e r a l s t a b i l i t y c h a r a c t e r is t i c s

of model I V.

NACA Langley Field

Va.

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Documents

19710066227