8/11/2019 Fin_Flare

1/9

THE APPLICATION O F A FIN-FLARE COMBINATION

FOR STABILIZING THE UPPER STAGE OF A

T W O ~ T A G EVEHICLE

JohnN. Lewak

Design Engineer

VeGicle

Design Department

Rocket and Space Division

Bristo l Aerospace Limited

Abstract

The Black Brant IV two-stage sounding rocket

wa s designed to meet scientific experimental study

requ irements at altitudes of approximately 1000

K .M.the second stage mus t be adequately stabilized

throughout a flight environment which includes

Mach numbers from 4 to 12, low dynamic press -

ures , and a relatively long critical stage separa-

tion period.

analyses indicated that fin stabilizing surfaces

would not provide sufficient stabil ity. At the same

time, the use of a conical flar e stabilizer was in-

vestigated and shown to be advantageous. Thus,the concept of a flare stabilized second stage was

pu rsued. Short ly before the in it ia l Black Br an t

firings, a wind tunnel study wa s carried out by the

Canadian National Aeronautics Establi shment ,

primarily to ve ri fy the theore ti ca l stabil ity analy-

s i s , which had been done fo r ze ro angle of a ttack.

They also indicated that non-linear effects could

prom ote stat ic instabil ity at smal l angles of attack.

The solution to this problem was found to be f in-

fl ar e combination which would provide the ne cess -

ar y stability throughout the operating Mach number

range.

rations resulted in an optimum design which may

now be considered fo r flight testing.

To satisfactorily perform these missions,

Ear ly i n the development phase ,

Subsequent tests on two fin-flare configu-

1. Introduction

The concept of a two-stage high performance

sounding rocket was propo sed a s one of a family

of sounding rockets to probe altitudes fr om 100 to

1000k . m .

Na tu ra lly, this vehicle wa s intended

for the upper end of this altitude ran ge , but, to

incorporate flexibility into the system, the second

stage would originate a s a single stage for employ-

ment at lower altitudes. The booste r, developed

initially as apropulsion test vehicl e, had al re ady

been flight proven in a single stage version.Ther efor e, while some modifications were neces-

sa ry to perm it mating of the two stages, the

motors and much of the hard ware for thi s vehicle

have been developed elsewhere.

2. Development of the Conical Flare Stabilizer

In the course of the prel iminary design of

the Black Brant

stabilizing assem bl fes were examined.

volved a three-fin configuration similar to that

used on the boost er, the other a conical fl are.

sustainer , two different

One in-

The initial analyses were car rie d out on bothasse mbl ies simultaneously and indicated the

de-

sireabi lity of using the fin assembly for the single

stage vehicle and the fl are on the two-stage

vehicle . The decision was made to follow-up this

approach and consider the flare as the only possi-

ble st ab il izer for the sust aine r. As complicat ions

aro se fr om this decision at the ,termina l stages of

the design progr am, it would be interesting to se e

how and why the flar e was sel ected as the s t ab i i

lizing unit.

To begin, a number of requirements wereestablished to govern the design. Pri ma ril y, the

philosophy of pe rformance and re li ab il ity at min-imum cost wa s adopted. Reliability was achieved

by using what would be existing hard war e, thus

reducing developmental cos ts. In addition, the

design and vehicle operation wer e to be kept as

simple asposs ib le , such that material costs

would be low and no complicated apparatus would

be needed for the vehicle to perf orm its function.

Fr om this philosophy evolved two particularly

important design requirements:

1) the connection between the two vehicles

would be broken by drag separation tech-

niques.

2 ) the sta tic margin of the upper s tage was

never to be le ss than 1 calibre (or1

second stage body diame ter) .

These requirements influence the design ofthe second stage considerably.

To conform to the procedure adopted during

the initial design phase, le t us examine the aer o-

dynamic properties of fin and flare stabilizers.

A typical example of the fin stabil izer asoemblyaerodynamic characteristics is shown in figure

1

for a 3-

fin sys tem that would be considered for useon this sus tainer. While the derivat ive of the lif t

coefficient presented has been referenced to the

body cross sect ional a rea and factor ed to include

body effects

Mach number, a characte ristic of fin stabilizers,i s clearly illust rated. Also shown is the predom-

inant influence of the lift coefficient on the stabil-

izing moment,

to be inve rsely proportional to the Mach number.

This decrement in lift coefficient and moment issaid to be destabilizing the vehicle, as il lustrat ed

the strong inverse dependancy on

which shows a distinct tendency

143

8/11/2019 Fin_Flare

2/9

in figure 2. Presented here ar e the total vehicle

aerodynamic characterist ics, the lift coefficient

derivative, and the centre of pressure, which are

determined by resolution of the nose and tail com-

ponents. The nose in this case is a cone cylinderwhose lift coefficient and centre of pres su re

location inc rease with Mach number. Thesefeatures are also destabilizing, but the tail con-

tributions are by far the determining factor of the

li ft and sta tic stability of the vehicle. Note that

the l i f t coefficient again displays a tendency to the

inverse Mach number proportionality and that thistrait has also been developed in the centre of

pressure location, a s it moves forward with in-creasing Mach number.

If the minimum -payload centre of gravityversus Mach number curve i s now superimposed,the deficiency of fins in stabilizing a vehicle which

has this range of flight Mach numbers i s very

apparent. Pr io r to the sustainer achieving a Mach

number of 6.5, the centre of pr es su re may be

seen to lie behind the centre of gravity, giving

ris e to apositive static margin and a statically

stable miss ile. At Mach 6.5, the centre of pres-

sure and cent re of gravity are coincident and the

vehicle i s neutrally stable. At Mach numbers

above 6.5, the centre of pressure li es forward of

the centre of gravity; the static margin i s negative,

and the vehicle is statically unstable.

Technically, it ispossible to side-step thiscondition either by including a control system to

provide ar ti ficial stability, or by imparting a

finite spin rate and thus gyroscopically stabilizingthe vehicle in much the same way as abullet isstabilized. However, in designing this vehicle,

neither method is appropriate. A control system

i s a luxury which the low-cost philosopy does notpe rmit , while a spun vehicle may be detrimental

with respect to user requirements.cas e must consider, therefore, anunguided, non-rolling vehicle fo r which the condition of sta tic in-

stability i s intolerable.

vehicle experienced an angle of attack during thisperiod, the fo rc e and moment systems which result

would destabil ize the vehicle more tending to

cause an angular growth. Eventually, the angle ofattack becomes sufficiently large to re sult in

vehicle break-up or, failing this, to prevent

successful completion of the mission.

then, the three-fin stabilizer can notbe used.

The design

If for some reason, the

Obviously,

The above discussion does not provide allow-ance for any fin stabilizer assembly other than

that originally considered, a three fin assembly

which, incidentally, is the current productionmodel on the single-stage version of the sustainer.

It i s possible to achieve the s tatic stability simply

by increasing the fin area, or even adding anotherfin. In this case, however, at the maximum Mach

number, the increase in weight at the aft end,

causing a centre of gravity shift, offsets anyapparent benefit of shifting the centre of pr es su re

further back, and stability may not be achieved by

this method. In other situations, these additions

might be the solution, but, before this method of

over-coming stat ic instability is adopted, theweight and drag penalt ies on performance shouldbe evaluated and consideration given to additionalmanufacturing costs in modifying the support

structure.

Returning to the problem at hand, it must berecalled that a conical flare stabilizer was eval-uated simultaneously with the fin assembly. The

fi rs t stabilizer considered had a terminal dia-

meter equal to that of the booster thus providinga smooth transistion in the inters tage region. It

was found to be inadequate, but, by increasingthe flare angle and, thus, increasing the terminaldiameter, a suitable fl ar e was found.

acteristic s of this stabilizer ar e herein presented.

The char-

Theoretically, the conical flar e behaves

aerodynamically like a cone, and the methodsoutlined here show that cone theory may be usedto predict the characteristics to a reasonableaccuracy. The fla re may be considered to con-

si st of a large cone from which a small cone,

having the same semi-

vertex angle and a co-

linear axis of revolution, i s removed.

this represents the procedure for determiningthe aerodynamic characteristics of the flare.

Since the semi-vertex angle i s the same, the l i f t

coefficient, referenced to the base are a of each

cone, i s also the same for a given Mach number.

However, when the coefficients are referenced to

the same cross-sectional a rea (general practiceis to use the maximum body cross-sect ion which

would be equivalent to that at the terminat ion ofthe small cone), the lifting power of the la rger

cone takes on a tr ue r perspective with respect tothe smaller.

generally

a function of the cone length, and, there-

fore, each may be evaluated in the usual manner.

The flare characteristics ar e, then, the resultant

coefficient and centre of pressure determined byremoving the small cone from the large , i n the

s a m e manner as the total vehicles characteristi cs

are calculated. It has been determined experi-mentally that attached flows about the flare wil l

cause a loss in l i f t effectiveness ranging from 0-

10 of the truncated cone value, depending uponthe nose shape and flow regime encountered. In

supersonic flow, for a configuration such a s thesustainer with its relatively slender nose, the

attached flow condition prevai ls, such that for de-

sign studies, the truncated cone values were

factored by 0.9.sented i n figure 3, and indicate much be tter

stability characteristics than the fin stabilizer

especially in the lift moment which now increaseswith Mach number instead of the rad ica l decrease

previously shown.

Basically,

The centre of pre ssure location i s

The analytical resu lts ar e pre-

Combining the flare Characteristics with the

nose-body propert ies and comparing the resultantto the fin-stabilized sustainer, the effectiveness

of utilizing a conical stabilizer may be seen

144

8/11/2019 Fin_Flare

3/9

(figure 4).this configuration is higher and the centre of

pressure is further aft.of gravity location, it can be shown that staticstability is maintained throughout this Mach num-

ber range, the minimum static margin being about1.3 calibres at approximately Mach 4. Since the

dynamic pre ssu re i s relatively low at this time,

aeroelast ic effects ar e negligible, and only rigid

body characteris tics a r e considered.

Above Mach 5, the lift coefficient for

Superimposing the centre

The results of figure 4, represent the terminal

design point in the development phase of the fl ar e

stabilizer . As this design met the stati c stability

requirements, it was adopted for use in the flighttesting phase. In this form, the fla re has two de-

trimental aspects. The more serious may be seen

from figure 4, which shows the minimum stability

occurs at the lower Mach numbers, and, hence, at

separation of the two vehicles.this wi l lbe discussed later.

that of the drag incurred by use of the fl ar e. Infigure 5, which shows the present flight configura-

tion, it may be seen that the flare extends beyond

the circumference of the booster. Therefore,

during the boosted portion of the flight, the fl ar ewill contribute to the base drag of the vehicle,This contribution, in relation to the total vehicle

drag, is quite small and does not influence per-

formance to any great extent in this phase of the

flight. Because the terminal diameter of thenozzle is only slightly smaller than the flare,

during the sustainer powered portion of the flight,the base drag contribution again is relatively small.

However, during the uapowered sustainer flight,

the base drag of the f la re contributes 40 - 50% ofthe total second stage drag. Fortunately, the un-

powered sustainer flight for this vehicle does notcommence until approximately 45 k.m. (150,000

ft.) , at which point the density is sufficiently smallas to prevent an excessive drag build-up. A drag

penalty of this nature could not be tolerated on thesingle stage version, which burns out at a much

lower altitude, and, consequently, the flare stab-ilizer is limited for use only on the two stagevehicle.

The implications of

The second factor is

3. Wind Tunnel Test and the Development of theFin F la re Combination

Recalling the facts that the minimum stability

of the fla re stabilized vehicle occurred at thelowest Mach number, and that the lowest Mach

number during sustainer atmospheric flight oc-

curred immediately pri or to second stage ignition,it may be seen that it is desirable to maintain theMach number as high as possible for this event.Due to the nature of the separation system, themost feasible time is shortly after the two vehiclesdisengage and the sustainer has preceeded the

booster sufficiently a s to prevent exhaust deflec-

tions. At this time, the Mach number has de-

creased to the Mach 4 region where stability i s

relatively low and the vehicle very sensitive to

atmospheric disturbances.

As the separation phase is extremely crucial

with respect to the whole flight, it was decided toverify the theoretical stability calculations employ-

ing a se ri es of wind tunnel test s.

made with the co-operation of the High Speed

Aerodynamics Section of the Canadian NationalAeronautics Establishment, who performed the

test s in their 1.5 x 1.5 m . (5 x 5 f t . ) blowdown

supersonic wind tunnel. The first ser ies of test s

on the fla re stabilized vehicle were performed at

Machs 3.5 and 4.5, thus covering the expectedrange of Mach numbers at separation. A theoret-

ical analysis indicated that at these Mach numbers

there would be no shock-boundary layer in ter-action causing separated flow.

photographs could not be taken to confirm this as-

pect of the study, the reduced data indicated thatunseparated Bow probably existed. Scatter in the

data reduced the effectiveness of the tes ts but two

important resu lts were obtained.

These were

While schlieren

1) the l i f t coefficient of the f la re was .only80 of the truncated cone values as com-

pared to the 90 used in theoretical

analysis, while the centre of pre ssur elocation agreed favourably with the theo-

retical estimation.

2) neutral Stability and subsequent static in-

stability occurred a t smal l angles of

attack when the reduced wind tunnel datawas applied to a typical flight case in-

volving a minimum net payload weight.

To reduce the effects of data scatter and to

verify these resul ts, a second ser ie s of tests wasrun at Mach 4.25 ,with apart icular emphasis on

low angles of attack.ed and when applied to flight cases indicated that

neutral stability occurred at an angle of attack oft 2.5O for a 18.1 k,g. (401b.) minimum net payload,

and at

payload, the Wference being attr ibuted to the

heavier payload, figure 6

Similar results were achiev-

for a 45.4 k.g. (100 lb.) maMmum net

It should be noted that the omission of a studyinto non-linear effects on veliicle stability does not

constitute an er ror in the fundamental design pro-

cedure. At that time, lit tle was known about flarestabilized vehicles, part icularly at angles of attack.

The result s of these tests do not increase the know-

ledge in this ar ea except in reference to this part-icular vehicle.

need for fur ther understanding of the aerodynamicsof flares and the implications of their employment

for stabilizers.

They do, however ,point out the

As stage separation took place a t altitudts

wherein high wind shea rs due to jet str eams exist-

ed, the low angle of attack requirements for stab-

ilrty could not be tolerated, since this implied

severe wind restrictions. To increase the stabihity

of the second stage ignition, it was decided to add

small fins to the stabilizer to utilize their low Mach

number lift effectiveness. Again, this was a yir

145

8/11/2019 Fin_Flare

4/9

tually an unknown region of aer ody namics , par t -

iculary at this Mach number level. As yet, t her e

is stil l no theory to adequately pred ict the charac-

te ri st ic s of this configuration. The only recourse

was, in wind tunnel testing, and fo r this study two

fin s planforms were examined, figure 7.

The results of the se tes ts , with resp ect to the

stabilitya f

the second stage a re shown in figure 8

for a Mach number of 4.25, and rol l angles of Oo

45O5 and 90.pe ct s to be found in these results.

There ar e s everal important as-

1) decreasing the sweep angle, thus increas-

ing the planform ar ea , increases the

static stability throughout the angles of

attack.

The forward centre of pr es su re shift at

angles of attack is not eliminated, but,

since the static margin at zero angle of

attack has been increased with the addition

of the fins , t he angle of at tack envelope

has 'now been extended to a tolerable

region.

2)

3) T h e stabi lit y of this configuration is roll-

angle dependent fo r any angle of attack

other than zero.

This last feature isprimarily t he result of

body shed vort ice s in teract ing with fin s causing a

loss in l i f t effectiveness, and may be seen fr om

the assyrnet r ical centre of pressure shift. In thecase of the Oo rol l angle, two fins ar e masked by

the body in the negative angle of attack region,

causing a mor e rapid forward centre of pres sur e

shift than for positive angles of attack. At roll

angles of 45O and 90, there is less interferenceand consequently more symmetrical shift.

these results , it may be seen that rolling the

vehicle, and thus continuously changing the rol l

orientation angle, would be beneficial fro m a

stability standpoint.

From

By interpolating the re su lts fro m two sets of

fins, a final, optimized planform maybe achieved.

This was done, resulting in a Configuration with

characteristics as shown in figure 9.

5. Flight Results

To date, the fin-

flare stabilizer hasbeen de-signed, manufactured, and structur ally test ed,

but no flight test ing of this configuration has been

made.

significant inc rea se in second stage dra g and

weight, resulting in loss in performan ce which has

been roughly es tima ted asbeingabaut 10

of

apogee altitude.

icle must suffer a n angle of a tta ck build-upsufficient to involve the nonlinear effects.

cause of such abuild-up must neces saril y be due

to a la rg e separat ion induce yaw, a n d l o r an en-

counter with a lar ge square wave wind she ar in

The addition of fins naturally rep res ent s a

To necessitate its use, the veh-

The

the order of 78 meters per sec .

separation assem bly design would be modified to

eliminate any separation distrubances. Wind

studies (ref . 8) have indicated the maximum

meas ured wind sh ear s ar e in the ord er of 24.4

meters per sec . over a 305 met er layer much

les s s eve re than the squa re wave magnitude re-

quired toproduce non-linearities.

wa s decided to investigate vehicle flight character-

istics using the d i n n e d stabilizer assembly, r e-

taining the fin-flare combination in reserve.sequent flight tes ts revealed the absence of any

separation disturbances, while wind-shears ap-

proaching the maximum non-linear design

case were not encountered.

tunnel test r esul ts indicate that it is desirable to

employ a fin-flare combination, it s need has yetto be demonstrated. However, flights with min-

imum payload (18.1 k.g. o r 40 lb.) m ay yet dict-ate the use of the

n e v

component, and, since its

ae'rodynamic

and st ruc tur al development is com-

pl et e, its incorporation into the BIack Brant Vsystem may be expedited without delay.

If

necessary, the

Therefore, it

Sub-

Thus, while wind

6. References

1. Ellinwood, J. W . , Parsons, W. D., and

Nakagawa, T . T . , Effectiveness of Fl ar ed

Afterbodies in Lift: Data Survey and Applica-

tion to the Exos Third Stage.

Report106-R2,

May, 1962. (CONFIDENTIAL)

Space-General

2.

Black Brant IV, Engineering Report No. 4132

Issue A, Bristol Aerospace Ltd. 1962.

4. Dixon R.C. , and Galway, R.D. ,Addendum to5 x 5 0001 and / 0002, National Aeronautical

Extablishment High Speed AerodynamicsSection, Wind Tunnel Data Report 5 x

5/0004,

Mar ch-April , 1964.

Dixon, R.C., Stability of Black Brant IV

Second Stage with Flare-Mounted Fins, National

Aeronautical Est abli shment, High Speed Aero-

dynamics Sect ion, Wind Tunnel Data Report

5 x 5:0006, September-October, 1964.

5.

7. Hayes, C . and Fournier, R.H.,Effect of Fin -

Fl ar e Combinations on the Aerodynamic

Chara cteri stic s of a Body at Mach Plumbers

1 . 6 1 and 2.20, NASA TN D-2623. February,

1965.

8. Sissenwine, Norman, Windspeed Prof ile , Wind-

shea r, and Gusts forDessgn

of Guidance

Systems for Vertical Rising Air Vehicles,

Air Fo rce Surveys in Geophysics, Special

Pro jec ts Laboratory Geophysics Rese arch

Directorate, Air F orce Cambridge Research

Centre, Air Research andDevelopmei9

Command, November, 1954.

146

8/11/2019 Fin_Flare

5/9

3 4 5 6 7 8 9 IO I 2 1 3

MACH NUMBER

X VERSUS

2095

3 4 5 6 7 8 9 1 0 1 1 1 2 1 3 3 4 5 6 7 8 9 IO\

12 I

MACH NUMBER MACH NUMBER

FIGURE

AERODYNAMIC CHARACTERISTICS OF FIN STABILIZER ASSEMBLY

MACH NUMBER

CENTRE OF PRESSURE

ANDCENTRE OF GRAVITY

LOCATlO NSVERSUS

MACH NUMBER

NET PAYLOAD WEIGHT -18.1Kg.d 25.4cm.

X

B

CENTRE OF PRESSURE

3 4 5 6 7 8 9 IO I 2 1 3

MACH NUMBER

L

FIGURE 2

AERODYNAMIC CHARACTERISTICS OF FIN STABILIZED VEHICLE

147

8/11/2019 Fin_Flare

6/9

3.54

3 4 5 6 7 8 9 IO 1 2 1 9MACH NUMBER

3 4 56 7 8 9

1011 I 2 1 3MACH NUMBER

FIGURE3

AERODYNAMIC CHARACTERISTICS OF FLARE STAB1MZER ASSEMBLY

3 4 5 6 7 8 9 IO I1 12 IMACH NUMBER

CENTRE OF PRESSURE

MACH NUMBER

3 4 5 6 7 8 9 1 0 1 1 1 2 1 3MACH NUMBER

FIGURE 4

AERODYNAMIC CHARACTERISTICS OF FLARE STABILIZED VEHICLE

148

8/11/2019 Fin_Flare

7/9

OVERALL VEHICLE LENGTH - l14.6m.OVERALL SUSTAINER LENGTH- 55.9m.

62.6cm.

FIGURE 5

BLACK BRANT IY FLIGHT CONFIGURATION

I

d

r---

c g

(45.4 Kg. NET PAYLOAD)

+n

IXcg

(18.1 Kg. NET PAYLOAD)w

. u

[Y

g-WIND TUNNEL2 I

I

I

I

13

5

O

II

I

I

I

I

I

Q

I

Y

i

2

4 I

1

0

II

LL I

3

I ITHEORETICALI

I

1

1

1

I

a

41 I

d =25 . 4 cm.

FIGURE 6CENTRE OF PRESSURE OF F L A R E STABILIZEDVEHICLE VERSUS A NGL E OF ATTACK

M = 4.25

149

8/11/2019 Fin_Flare

8/9

FIGURE 7FIN- FLARE COMBINATIONS-TEST MODEL CONFIGURATION

I

I

LOOK INC

FORWARD

vXcg l8.l

Kg. N E

LEGENDROLLANGLE I HALF SPAN FINS I FULL SPAN FINS

O0 Q 6 6

450

90

r PAYLOAD d=25.4cm.

FIGURE

CENTRE OF PRESSURE OF FIN-FLARE STABILIZEDVEHICLE

VERSUS ANGLE OF ATTACK ATM=4 25

8/11/2019 Fin_Flare

9/9

CENTRE OF PRESSUREVERSUS

ANGLE OF ATTACK

STRUCTURAL LIMIT

X c g 18.1

Kg. NET PAYLOAD)

FIN DIMENSIONS

FIGURE 9

OPTIMIZED FIN-FLARE CONFI ;URATION