AERODYNAMIC FEATURES OF HYPERSONIC SPEEDS

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• Collection: General CIA Records
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• Release Decision: RIPPUB
• Original Classification: U
• Document Page Count: 12
• Document Creation Date: December 23, 2016
• Document Release Date: December 24, 2013
• Sequence Number: 24
• Publication Date: September 2, 1958
• Content Type: REPORT

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DATE Of REPOqT
k'
2 Sept 1958
SUBJECT (Descriptive title. Use individual reports 101- ..14,?111???
AERODYNAMIC FEATURES aF HXPERSONIC SPEEDS
.s
SUMMARY (Give summary which highlights the salient factors of niriativif reptt. Begin narrative text on Al Form I12a
unless report can be fully lasted on Al Form 112.-Liat inclosures. including number of copies)
Forwarded herewith is a translation of an artic14, entitled "Aerodynamic
Features of Hypersonic Speeds" (Osobennosti aerodinamiki giparsvukavykh
skorostey), written by Docent, Candidate of Technical Soiences-M. L. Gorman
and published in P: Vestnik Vozdushnogo !Iota (The Herald of the Air Fleet),
No. 11, 1957, pp. 56-64.
The article describes some aerodynamic features of hypersonic speeds and how
a hypersonic flight is investigated in wind tunnels and impact tubes by means
of models.
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1. Fig. 1 - Schematic diagram of ,the flow around a plate at high Ma
Fig. 2 - Diagrams representing eseffiAien$is of the lift force of a
plans plate for differimti-
Fig. 3 - Diagrams representing 1;04'017 06i:double-wedge, triangular
and convex profiles, caledlited fbi different parameters of
similmrity K.
2. Fig.. 4 - Schematic diagram of separation of the flow at the trailing
edge of a supersortc airfoil.
Fig. 5 -. Diagrams representing the distribution of pressure over a
convex airfoil.
.Fig. 6 - Schematic diegram showing three regions of hypersonic flow
around a plane plate.
-Fig. 7 - Diagrams showing the dependenci..Of-imiaikm preiteure coefficient
of a plane plate on M.
Fig. $ - Diagrams showing the dependence of the "distance of interaction?
on Re and M.
3. Fig. 9 - Diagrams shoving the temperature of retardation for bodies
moving at different Nh and different altitudes.
Fig. 10- Schematic drawing of a hypersonic impact pipe.
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unauthorised person is prohibited by law It may not be reproduced in whole or in part, by what. than United States Air Force
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3 9
THE FEATURES ar HYPERSONIC spI9 AEAODYNA.MrdS
A continuous progress of supersonic aviation presents a seriee of new pro-
blems associated with the flight of flying devaces at high MS, .For in-
stance, to be sure that the range of the rocket will reach 10,000 km, the
speed must correspond to Mil 20 (Ne are bamedon the sonio speed in strato-
sphere). For an Earth's satellite M should be about 25. It uhould be re-
minded that the meteors travel at M equal to from 30 to 100. Thus, because
of the development of missiles, rockets, and the possibility of interplane-
tary communication the aerodynamics of hypersonic speeds acquire greater
and greater importance.
A hypersonic flow possesses,a series of characteristic features which appear
at high Nis. Ta explain them,. one4shqgld investigate the physical picture of
the streartlinWeItypersonic now around a body. It is known that the angle
of inclinAtion.of the shock and the angle of disturbances decrease with the
rise of M. Furthermore, at high Ms these angles are so small that the shocks
as well as the lines of disturbances strive to adjoin the surface of the body
over which the flow proceeds. It is possible to say that the shack becomes
almost parallel to- the direction of notion. In this ease the region of dis-
turbances is very small and can be compared with the region occupied by the
boundary layer.
For instance, if at Mac 5 the angle of disturbances for a thin body is equal
to 11.54?, then at Nix 10 this angle will be equal to 5.70. The calculations
show that at M = 10 the angle of inclination of the shock for a double-wedge
airfoil amounts to 8.2?.
In Fig; 1 is shown the flow over a plate at very high M. Since the increase
of pressure outside the shook at high M by far exceeds the decrease of pres-
sure in the rarefied flow, one can approximately assume that over the upper
surface of the plate is a vacuum. On the lower surface of the plata an-
oblique shock adjoins the plate very closely. Therefore the flow sisown in
Fig. la excellently correspondswith the flow about a plane plate shown in
Fig. lb, which proceeds according to the Newton's theory of impact. Thus,
it was found that the Newton's theory which yields wrong results at low
speeds, fits well for very high Ms. Presently, the so-called linear theory,
which is based oil the assumption that bodies (wing fuselage) of-small thick-
nesses at small angle of attack are treated, is widely used for practical
problems of supersonic aerodynamics. In such ease the existing shock can be
-replaced by a line of disturbances, and thus simplify the exact theory. For
instance, according to this theory the pressure coefficieho Changes in ?direct
proportion to the angle of attack.
For a hypersonic flow the linear theory becomes uselese.
planation of this can be given: this theory can be appliiit _
maximum angle at which the flow is declined due to the presence 0:tne- 9
is small in comparison with the angle of disturbance of the free-Stream. But
in a hypersonic flow the cone of disturbances is so 'set that their A.4)e can
be compared with the angle of inclination of the plate's nose. _
X
This leads to a fact that the coefficient of the air pressure-becomes pro-
portional to the square of the local angle of attack. Hence, the linear
theory appears to be approximately correct only for small angles of attack,
up to Ms between 4 and 5t From Fig. 2 it is seen that already at M higher
WARNING This document contains information nffecting the national defense of the United States within the m
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than 3 the linearity of the graphs, representing the coefficients of the
lift foroe of a plane plate, is distorted. In addition, it should be born
in mind that according to the linear theory the aerodynamical coefficients
do not depend on the nature of the gas and the character of the thermo-
dynamic process. However, at higher Ms the nature of the moving gas changes?
Due to a strong impact during the collision of molecules traveling at high
_speeds, their compound atoms start to vibrate respectively to their;tan
position. Additional degrees of freedom will appear,W41,,iiir-=,
to an increase of'speeific heat at constant volume and toa itecnviii of the
ratio K, from K =L.-20:40*1-45HIA
As to the term "hypersonic" flow itself, the results of the linear theory
alley to set the condi-gone' limits of this flow? At M greater than 5 it
can be assumed that v144- 1111.14 (for instance, at PI= 5 P4 1 =
This leads to a simplification of some known formulas of the lift force and
the head resistance of a double-wedge airfoil. It can be assumed that the
term Impersonic" concerns Ms higher than 50
The following law of similarity of hypersonic flow can be defined. If
bodies having similar forms, but a different relative thickness 6, are
placed into a flow of a different M so that the parameter NE will remain
constant, then the flow will be also similar, i. e., the streamlines will
be alike. This means that if a hypersonic flow is over the airfoils whose
thickness and curvature are distributed identically while the angle of
attack is proportional to the relative-thicknesal -pen their aerodynamic
coefficients depend only on the criterion Onamilaiity NE.
To illustrate this law, an example is given.
Suppose that coefficients cy and cx of a symmetric airfoil with a thickness
al . 10% at an angle of attack Ocs 5?, tested at M = 5, are known. The
qtestion is: what will be the aerodynamic coefficients of a geometrically
similar airfoil whose relative thickness B, = 5%, and to which Mwill they
correspond. Since there is a similarity, The parameters ofsimilarity
should be equal, i. e., M02. It follows that it = 10; in addition,
the similarity of the airfoils will be at c44, 4, U2 p 1. e., 0(= 2.5?.
ati ? - -EI--
From the formulas for the coefficients it follows that
E2 N?
C2_-(_2 3 2
- 0.25; cx2
1 0.125.
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0 31
Yi- 1 ex1
Thus, the criteri4e-efAi1i!s,rity by the aid of the known coefficients of
one airfoil pei4t410-01174sily to determine the coefficients for similar
airfoils. For4iAllaitiitm, polars of variousairfoils, computed for three
values of similarity parameter, are given in Fig. 3.
Although the hypersonic theory of similarity is only an approximate one, it,
nevertheless, corresponds well with the experiment and allows to compute
the aerodynamic characteristics of the wings and the bodies of revolution?
?
An important distinctive feature of the hypersonic flow is the strong effect
of viscosity which leads to an essential interaction between the boundary ,
layer and the shock wave?
WARNIN6 ? This document contains information affecting the national defense of the United States within the meaning of the ?
Espionage Laws, Title 18, U S. C., Sections 793 and 794. Its transmission or the revelation of its contents in any manner to an
unauthorized person is prohibited by law It may not be reproduced in whole or in part, by other than United States Air Force
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DuringAhe invee eation of a supersonic flow, a general approaCh can be
aaliedith r a.rd to the analysis of viszsity. We assume that the effect
of vimpoele- concentrates in a thin layer joining the surface: of the body,
the soleleed bouelary layer, and that it can be determined as i correction,'
to a notlyiscous flow. At small supersonic 'speeds the boundary layer and the
ehock-wails in the forward peet of the airfoil are located-far fr.* each
Other, co that the effect of their interaction can be neglected: It is
true that, in thi. case, the effect of the boundary layer shows up near the
trailing edge, ainc,e in this region the shock wave and the boundary layer
may big the cause et the separation of the flow. This occurs because the
incrbiteed pressure behind the tail shock P2, being unable to propagate
forward (at-Wpersenic speeds the disturbances do not spread forward),
? penItrates into region BDC (Fig. 4) through the subsonic part of_the-
boundary layer. 'raised pressure in this region causes a zeParkion
which prevents -a further expansion of the flowe and thus, the preemie.? in
the region BDC appears to be greater than the theoretical calculations,
would show. This is clearly seen from the curves representing theedistri-
bution of the pressure. From Fig. 5 it is seen that the theory and the
v
experiment coincide well in the forward part of the airfoil; they differ
only in the rear part of the airfoil-in the region of an increased pressure
(decrease in rarefaction).
A different picture appears at hypersonic speeds. Firstly, the, boundary
layer behind a strong head shock wave over a very thin airfoil has a larger
thickness as compared to that of the body over which the flow proceeds.
The fact is that the shock wave at high supersonic speeds adjoins the sur-
face of the airfoil so closely that the entire region between the surface
and the sheck vave should be considered as a region of viscous flow. Con-
sequently, when the parameters of a shock wave are being determined, the
effect of viscosity should be considered, and when the friction on the sur-
face of the plate iv being evaluated, the-effect of the shock must be es-
pecially taken into aceoeut.
Thus the flow in a bound:el, layer strongly affects the stream in the regien
between the outer eurface of the boundary layer and the shock wave. For
instancepempethe nose of he plate the boundary layer has a big curvature
140I0hfith;404714.!- bends the heAd &lock. Disregard of viscosity would lead to
ifSli-jAteliti3O?13?here
the shock :1011L1 appear as rectilinear. Because of this,
the 'region'of the boundary layer and the region between. the 'Shock and the
outer surface of the bogneary layer cannot be investigated separately, since
their interaction must be ..aken into account.
If, for insi:ance, one will examine the flow around a plane plate at hyper-
sonic speed, then the entire field of the flow, disturbed because of the
presence of the plate, can be divided into three regions (Fig. 6).
Close to the plate is a boundary layer in which the effect of viscosity is
strong. eeee-
?-?-zte "4-r;
From the nose of_thet:p? ? eaepowerful shock wave appears; at
'
the nose of the pike t Ma on of the shock is big, afterwards it
sharply decreases and, for_instance, at high 14 the shock wave rooks like
drifting along the plate. A region of undisturbed stream is above the
shock wave.
- - - - ? -
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Finally, between the shock wave and the outer border of the boundary layer
is a small region of the flow in which the effect of viscosity can be
neglected; across this flow a big change in pressure and density can be
observed.
Owing to that the boundary layer thickens, the shock wave at the leading
edge has a curvilinear form. As a result of this, the pressure along the
plate-is not constant and can be higher than that of the surroundings.
Furthermore, this excess of pressure depends on M and Re (Reynolds num-
ber - a dimensionless magnitude defined by the ratio of the inertia forces
and the viscosity forces in the flow). For instance, at MI= 5.8 and a
distance of 0.63 am from the leading edge the pressure inCreases by 70%.
The increase of the thickness of the boundary layer can be characterized
by the following: at M 7 the thickness of the boundary layer is 10 times
greater than at MI= 1 at the same Re (Reynolds). At great, distances from
the leading edge the shock wave and the boundary layer drift apart by a
considerable distance so that statically the pressure does not change any
more.
It was already shown (Fig. 5) that the boundary layer at low supersonic
speeds (and also at subsonic) affects the distribution of pressure, first
of all, in the tail part of the body. At hypersonic speeds the thickening
of the boundary layer becomes so considerable that it provokes an essential
change in the distribution of pressure over the nose part of the airfoil.
At the same time the usual occurences of separation in the tai/ part of the
body may still happen. .
?e,
The rise of pressure leads to a situation that the surface friction becomes
higher than it would be according to the theory of the boundary layer in
supersonic flow. The re3istauce coefficient of the laminar friction Ef
decreases with the rise of NI when the interaction is not taken into account,
and it increases at hypersonic speeds when the interaction is taken into
account (Fig. 7)0
In addition, at low supersonic speeds (up to M = 3) the coefficient Uf
does not depend an Re, while at hypersonic speeds it essentially depends
on Re.
For characterization of the mutual interaction ott4eAboundary layer and the
shockwavetheideaenthedistance.of plTteiacen1Iiintroduced, which is
. ? ? ? - ? "
expressed as a ratio Xo where X0 is the distanoo-fion'the nose to the point
e: -4
where the ske4k*I,-e:rdeif4e4A1Tinte disturbance wave, and L is the chord
of the bodiqFig..65'. Imo in Fig. 8 it follows that for a given
Re the effect of the shock projoagate's farther downstream when M is 'higher*
A special importance acquires the effect of the temperature at hypersonic
speeds. A conekjerejblel*Otease of temperature along the head shock leads
to a large altellaatetranefer to the body over which the flow proceeds.
This also affects the tlow close to the surface of the body.
Thus, behind the shock and along the boundary layer the temperature may become
so high that the chemical features of the gas may change. This loads toa
breach in the static .equilibrium between different kinds of internal energy
of the gaso
WARNING. This document contains information affecting the national defense of the United States within the -meaning of the
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unauthorised person is prohibited by law It may not be reproduced in whole or in part, by other than United States Air Force
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.Indeed, if thetemerature of the gas after crossing the shock is sap:-
cient to cause sqme4Oditional degrees of freedom of molecules, then a
dissociation of gas).0ay happen, i0 e., a change in the chemical compound
will occur as a result of disintegration of its molecules into more sim-
ple molecules or atoms.
1 PAGES STAT
A still higher tmaperature may cause the ionization of molectiles. A
change in the'strmdture_pf gas =ate seen during the tests qf models on
ballistic installations:or by free firing. So; for instance,, by firing
a ball in Xetioniat 14 higher than 10 one could see the ionization of Xenon.
During another experiment, when a steel ball, 13.5 mm in diaMeter, was
fired into a reservoir filled with Freon at a speed corresponding to
M 9.27, the disintegration of Freon was clearly seen. It shows how
important is the-problemassoeiated with the investigation of the chemical
:Taa,turqs if gas near the shock waveo
The-effect of additional-degrees of freedom of intermolecular motion on the
features of gas depends on the temperature and the pressure and varies with
the altitude.
,
In Fig. 9 is shown the tempei100411-
at different altitudes. Theqiigram-of AtcAdoil gas (parabolic law of the
temperature rise with the rise of PO does not reflect the real features of
a real gas at high temperatures. So, for instance, at MI= 10 an ideal gas
shows a temperature of retardation equal to 4700? X, while a real gas de-
-pending on the altitude has considerably lower temperature of retardation.
;
Presently, several me ,
ing a hypersonic flow can be recommended.
To the first method belong a baiiiiiic installation and a free firing. Here
the models are tested on ballistic installations which catapult the models
into motionless air at high speeds?
Recently, the installations for free firing have been also used. As an
example, a cannon which can provide a speed up to 20,000 kilometers per hour
can be taken. The cannon may also fire into a tank in which the pressure
of the gas or the air is variable, thus creating the conditions of necessary
pressure or altitude. The Maw. can be also:Ado* by, a high-speed cannon
into a supersonic stream generated hie. spec*IIn this way the cannon
can fire the models in a free-.light at Nh oeifie7oider of 20 and moi4e.
Applying special vuovicifciT;;:.tk 7'.recording of time, it is possible to photo-
graph the models a1,54014,--iiiii041r one tenrillionth of a second. Although
such cannons create aoie 1 iineli of' a real flight and real temperature,
nevertheless measurement of aerodynamic
:-.tere4ik4n.ballistic
installations, is rather diai .. -i..4...zt
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2.4j5?.
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Striving to approach the real :ewe . ,.... _ . c ed aerodynamic or
?rocket nvehicle" had to be developed. This is a special railroad ,path of
3-7 km length with some apTiliary mechanisms. Vehicles of the present, de-
signs can reach a speed up to 1500 misec, while the acceleration rises lip to
1-
100 g.
WARNING This document contains information affecting the national defense of the United States within the meaning of the
Espionage Laws, Title 18, U. S C., Sections 793 and 794. Its transmission or the revelation of its contents in any manner to an
unauthorized person is prohibited by law It may not be reproduced in whole or in part, by other than United States Air Force
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Recently, investigating the hypersonic flow impact pipes were used, usually
designated for the study of explosions (primarily atomic explosions), since
the wind tunnels do not provide results comparable to those reached under
conditions of a free flight. The most importank problem associated with
a wind tunnel is the sharp drop of temperature between the receiver and
the working part, which is the result of the expansion of the gas at high M.
If the air is delivered at a normal temperature (for instande, at To = 3000K),
then the air temperature in the working part at M = 5 and M= 10 will de-
crease up to 50?K and 14.5% respectively. At an atmospheric pressure the
temperature of liquefaction of the air is 780K. It means that in cases
under consideration a condensation will occur. To prevent such phenomenon
the flow must be heated.
Even having a well developed heating system it is difficult to acquire a
temperature of retardation in the working part comparable to that generated
in a free flight.
An impact PiP941.sta ,e0sentially, of along channel closed at both ends
and divided by a dis& into 'tio19.*.iitiOnkiliiich contain gases at different
pressures. Wring arePiire of the diaphragm a straight shock appears which
propagates into that part of the pipe where the pressure is lower. The ex,-
pension of the compressed gas creates a region of a high speed flow. It is
obvious that a settled motion proceeds only during a short period of time
(1000 microseconds at a pressure difference equal to 2500 atm).
A steel cylinder (1) in Fig. 10, which is abre-,to-vithstanda pressure up to
from 150 to 200 atm, serves as a compression chambei-of-the.pipe. It is
connected with a chamber of constant expansion (3). Between the compression
and expansion chambers is a plastic diaphragm (2).
Along the section of constant expansion are two observation points in which
either piezoelectric transmitters or glass windows, to make photographs of
the flow, are arranged. The chamber of constant expansion is connected with
the divergent part of the pipe (4) in which the expansion of the flow beyond
the shock is going on? thus ensuring a high M in the working part.
In the working part (5) there are also windows for both the optical stud/
of the flow around the models, and for investigation of the pressure in the
flow Or over the model.
Different MS ereie,NniTed,hy changing the ratio of pressures in the sections
of compression aiad zar.f&ction,lan,d also by changing the degree of the
nozzle is diverge**. - '-
Hypersonic Na cawbe.reixed in the impact pipes by using a mixture of hydro-
gen. and heliumgas.The deficiency of impact pipes lies in
th-4iffieN4 . .serodyiimic forces which affect the model,
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Presently, hyperdanic wind tunnels are widely used for investigation of flow
at very high Na
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There n6 essential difference between theapersonio and hypersonic
tunne4 (for instance; at NE= 2 and NE= 10); fE both cases the divergent
noizie, whIre the necessary supersonic speed is reached, is beyond the
cr#icaLsee,tion. However, there are qualitative differences.
-jd
Inability to get sufficient power and necessary cooling is the basic diffi-
culty in dealing with the ordinary supersonic wind tunnels. The same diffi-
culties appear also when the hypersonic speeds have to be reached.
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The rsitin. cuUarity pt.-A-hypersonic tunnel is the variation of the,tempera-
ture and pressUreAepending on 141 which leads to a liquefaction of-the
coolWeir. This creates conditions in the flow which are different from
the natural ones: the liquefaction which affects the aerodynamic charac-
teristics, the riseof the inclination angle of the shock, the change of M.
To prevent the liquefaction of the air, the.. compressed air must be heated
if bore it goes to the working part. However, the compressed air cannot be
heated above 250?C since the ability of the steel reservoir to withstand
the pressure decreases with the rise of the temperature. Therefere, in a
wind tunnel, working with air, it is not possible to obtain M higher, than 10.
To het higher Ms, other gases must be used: nitrogen, hydrogen, helium, freon,
xenon.
A series of problems appear in connection wit!44)104,eliudeneirefhypersonic
flow. Low densities and temperatures increadeVery401m0gEthe-thipkness
of the boundary layer in the tunnel. beside the difficulty it selecting
the nozzle, a large thickness of the boundary layer, Which distorts the
posture of the shocks and angles of disturbances, makes impossible to detdr-
mine M of the flow by means of the shock's inclination or by the line of
disturbance.
For visual observations of the flow at high Ms the phenomenon of the
luminosity of nitrogen due to its electrification (oxygen or argon can be
also used but their luminosity is less bright)" is used.
The new branch of aerodynamics, the aerodynamics of hypersonic speeds, pre-
sently under development forms the theoretical basis of flight of inter-
continental and ballistic rockets and missiles at high supersonic speeds.
By investigating their characteristics one encounters a series of new and
interesting problems the majoiity of which require special methods.
STAT
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WARNING: This document contains information affecting the national defenstrOkthe,;,.?United States
Esititeri.lige Laws, Title 18, U.S. C., Sections 793 and 794. Its transmission or the reve.atiorf,of its con tAiiiiitnis".y..
una6i/iorized person is prohibited by law. It may not be reproduced in whole or in 'Part, b;r other than tl .41
Agericie's, except by permission of. the Director-of Intelligence, USAF A
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Fig. 1 - Schematic diagram Of the. flow
around a plate at high 14.
PunetIm 51runepleyfr
WAIF
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firms
II
Roams
Iv:1mm
fritiEWM
3)
2 8 V
fisoft omaltu -do
1) Expansion
2) Shock
3) Vacuum
Fig. 2 - Coefficients or the lir;
force of a plane plate for
different Ms.
1) Linear
2) gypersanic
3) Angle of attack
4) Uncampressible
et 80
60
S
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2? 40
'4
? 20
0
e 0
? 40
? 20
--. 0
N.,
40
20 ' ONININ?
0
0 60 qv cs
K.0.7
Aist
-NO
fl
?,/,0
210mmvumenvw(2
N0,414044ueffm
.0060800 cenpondlenum
Cs
_J
Fig. 3 - }tiers for a double-
wedge, triangular and
convex profiles, cal-
culated for different
parameters of similarity K.
1) Relative coefficient of
lift force.
2) Relative coefficient
of head resistance.
trayit -
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Cffavort
17otpom c.o? Il
Fig. 4 - Separation of the flow at -
the trailing 3dge of a super-
sonic airfoil.
1) Boundary layer
'2) Shock
3) Subsonic part of boundary
layer
Fig. 5 - Pressure distribution
over'a convex airfoil
whose-thickmess is 8.8%,
the angle of attack = 0
and M 2.13. On the dia-
gram can be seen the in-
creased pressure at the
trailing edge of the air-
foti.
Fig. 6 - Three regions of hypersonic
flow around a plane plate.
I - Boundary layer
II - Non-viscous flow
III - Free stream
1) Shock wave
2) '.4ve of disturbance
5 1 7 II V Ai
olueso Al
Fig. 7 - Dependence of median pressure
coefficient of a plans plate on
M with and without the considera-
tion cf interaction.
1) Resistance coefficient of fric-
tion
2) With consideration of inter-
action
3) Without consideration of
interaction
Fig. 8 Dependence of the "distance of intAr-
action" on Re and M
1) The distance of interaction
11810'31 -
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Fig. 9 - The temperature of retardation for bodies moving at different
Mb and different altitudes.
1) Temperature of retardation
2) Ideal gas
3) Real gas
4) Dissociation
5) Fluctuating temperature
6) M of flight based on the speed of sound in the stratosphere
Fig. 10 - Schematic drawing of hypersonic
1) Pressure chamber
2) Diaphragm
3) Section of Et constant expansion
4) Divergent section
5) Working part
Impact pipe.
6) Observation section
7) Counter
8) Amplifiers
9) Transmitters
10) Spark discharger
11 .?('7(.!!'1.
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