(SANITIZED)WORK ON MISSILES, 1936-1952

• Document Type: CREST
• Collection: General CIA Records
• Document Number (FOIA) /ESDN (CREST): CIA-RDP80-00810A004600160009-8
• Release Decision: RIPPUB
• Original Classification: S
• Document Page Count: 39
• Document Creation Date: December 20, 2016
• Document Release Date: August 10, 2007
• Sequence Number: 9
• Publication Date: September 21, 1954
• Content Type: REPORT

Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
INFORMATION REPORT
US OFFTC:iALS Oaly
41
RD
THE SOURCE EVALUATIONS IN THIS REPORT ARE DEFINITIVE.
THE APPRAISAL Of CONTENT IS TENTATIVE.
(FOR KEY SEE REVERSE)
Discovery of the Multiple-Shock Supersonic
Laws of the Two-Dimensional: Multiple-Shock
3. The Inverse Meier-Prandtl Flow.
4, Conical Flow.
a. Computation and Representation.
b. Limits of the Body Angle of a Cone in the Case of
Laminar Flow.
640592
Outline of Contents
it9ris niitl.ine is not exactly adhered to,but its points are covered in the report.)
25X1
This material eontatns Information affecting the Na-
tional Defenee of the United States within the mean-
Ing of the Lplonate Laws. Title 1$. U.S.C. Seel. 799
and 794. the tranmmiadon or revelation of which In
any manner to an unauthorised person to prohibited
by law.
REPORT I
II, T o t rohlo : The fl iffusor in i! erodynen.lic 5;nt;ine.
1II. The :ultiple-Shock Supersonic Diffusor. Theory.
NO. OF PAGES
REQUIREMENT NO
REFERENCES
 -2-
5. Three-Dimensional Rotation Symmetrical Aultiple-Shock :)J.ffucore.
6. Missiles without Impact Wave Resistance.
7. Three-Dimensional Multiple-Shock Supersonic Diffusors with
Variable Intake Cross Section.
IV. The Multiple-Shock Diffusor in Practical Research.
1. Experimental Arrangement and Experimental Results of Two-dimensional
Two-Shock Diffusors. Boundary Layer Exhaust.
2. Experimental Arrangement of Three-dimensional Multiple-Shock
Diffusors. Results of Experiment.
V. Diffusor and Nozzle in Reciprocal Action.
1. The Two-dimensional Meier-Prandtl Diffusor.
2. The Three-dimensional Meier-Prandtl Diffusor,
3. The Free (Without Wall) Supersonic Nozzle.
4. Meier-Prandtl Flow Combined as Diffusor Flow and Nozzle Flow for
a Nondissipative Change of Direction of a Supersonic Flow.
VI. Nozzle with Turbulent Flow.
1. Deriving the Equations of Motion.
2. Representing the Results.
3. Practical Consequences for Turbulence Nozzles.
4. Turbulence Nozzles in Turbines.
5. Turbulent Flow for the Production of a High Vacuum.
VII. Use of Supersonic Diffusors in Aerodynamic Engines. General Points
of View.
VIII. The Lorin Engine. Theory.
1. The Efficiency of the Lorin Engine with a Very Small Power Supply.
2. The Efficiency of the Lorin Engine with a Greater Power Supply.
3. The Efficiency of a Ram-Jet Missile.
4. The Efficiency of a Ram-Jet Missile at Greater Mach Numbers.
Average Efficiency. Range.
IX. The Operation of the Ram-Jet Missile. Physical-Chemical Problems.
1. Problems of Self-Ignition.
2. Problems of External Auto-Ignition.
3. Problems of the Rapid and Complete Reaction.
SECRET - US OFFICIALS ONLY
 Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
SECRET - 113 OFFSCTALS ONLY
-3-
The Problem: The Diffusor in Aerodynamic &gines.
In all aerodynamic engines the following processes take place indcpendently
or in combination with others:
a. Nozzle Processes: With a decrease in pressure there is an increase
in the velocity of the gaseous medium.
b. Diffusor Processes: The pressure is increased when the velocity
of the medium is decreased.
c. Mechanical energy is either supplied to or derived from the medium
by moving engine parts. It can always be proved whether such a
process is combined with a nozzle process or with a diffusor
process.
d. By means of combusion or a heat exchanger, energy in the form of
heat is either supplied to or derived from the medium.
 Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
The quality of at ^erodYnaic (:n?i.nc (;Ga)(.nd o u::
the above processes while ':ee;;in ; he cu .Tory i. for c.: e very :. c.ll.
a. Nozzle processes can be carried out in such a way that the entropy
increase is limited to an insignificant and unavoidable amount
caused by surface friction.
b. Diffusor processes have up until now been carried on with sufficiently
snail entropy increases only at subsonic velocities. No matter
how good the diffusor flow may be, the entropy iiicrea c is cl.eer~s
greater than in the case of nozzle flow. ll_i.ffusors which :;orl:
with an effi.ci?noy equa]. to tint of nozzles hr:?e been used :t sub-
sonic velocities only with special dissipative o;)craL"ons ('('r uiary
layer control, l ia,;nus effect),
In the case of supersonic flow, diffusors are not possible as reverse
Laval nozzles.
Diffusor as the
Reverse of a Nozzle
b.. In the case of supersonic flow.
a. In the case of subsonic flow, a diffusor as the reverse of a
nozzle is theoretically possible (assuming a careful :ilariu,:).
b. In the case of supersonic flow, a diffusor as the reverse of a
nozzle is theoretically impossible. A plane shock wave will
always be formed in front of the intake of a conically constricted.
tube. Shock waves are always produced when the attencr;. :Us made
to retard a supersonic floe. pressure increase by seaas of a
diffusor. With an increasing i,ach nwcioer the inefficiency, caused
by the entropy decrease in the plane shock wave always increases
very rapidly.
c. The energy exchange between moving engine parts and the flowing
gases as the power medium always shows a slight, unavoidable
amount of entropy increase as a result of surf:co friction. A
considerable limitation in the form of aerodynamic engines of all
types is imposed not only by the fact that the flow at moving
engine parts is accelerated with a decrease in pressure, but also
by the fact that there are always areas in which the flow is
retarded with an increase in pressure. The greater the retardation,
the steeper the pressure increase will be, and the greater the
losses and entropy increase will be as a result of flow separation,
shock waves, and turbulence. Thus, in designing aerodynamic
engines, and especially in the case of compressors, the attempt is
always made to distribute the pressure increase on the turbine
wheel and the diffusor so that a too rapid local increase of
pressure can be avoided. Up until now supersonic velocities in
 -5-
compressors have been avoided because of the great difficulties of
diffusor operation in supersonic flow. This limited considerably
the operational efficiency of compressors.
d. The processes of heat input and heat exchange are not directly
influenced by the problems connected with the diffusor, as far as
efficiency is concerned. On the other hand, if a solution to the
problems connected with the diffusor can be obtained so that aero-
dynamic engines can be operated with considerably higher pressures,
then much smaller heat exchangers can be built. In addition, the
processes of combustion will be able to be carried on with greater
efficiency at higher pressures, combustion chambers will be able
to be built smaller, and the dissociation of combustion products
in the case of some combustion reactions will be able to be
avoided through increased pressure.
This survey of the principle processes in aerodynamic engines is intended
to present and interpret the chief significance of the diffusor problem,
the transformation of any velocity into pressure.
Diffusor Processes in the Case of Supersonic Velocities
Subsonic diffusors have been so well studied, it appears that no appreci-
able progress in their development is to be expected in the foreseeable
future. To be sure, the processes in conically extended tubes need further
study in the range of the Mach numbers between 0.5 and a little below 1.0.
It appears that the most favorable angle of divergence with these Mach
numbers is smaller than the eight degree angle generally considered to
be ideal. In contrast to the well investigated processes in subsonic
di fnrsors, little besides the plane shock wave has become known about
dit(iiors used at supersonic velocities. Busemann and his co-workers in
Braunschweig from 1938 to 1942 attempted to explain, by theory and by
experiments in the wind tunnel, what possibilities exist for building
supersonic diffusors, and especially whether it is possible to force a
flow pattern which represents the reverse of the flow in a Laval nozzle.
Research'on these flows was facilitated by the process (hodographic
process) of Guderlei (Busemannis co-worker in Braunschweig in 1942),
which is a further development of Prandtlos and Busemann's process for
the production of two-dimensional supersonic flows.
Figure 2
Flow within a Short,
Convergent, Conical Tube
The supersonic flow coming from the left strikes at S at the edge of a
conical surface SB. The streamlines close to the edge are deflected
 Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
from the axis by the angle fl. The deflection of the streamlines propagates
along the conical surface SA of an oblique shock wave. Behind the oblique
shock wave the pressure has increased. If the course of flow is extended
from the edge S in the direction of the oblique shock wave SA, all the
..ar"?n+.ers of flrn,r hchinJ the shock, including thn shock angle 1/ r.n the
-/, i rnr'.i.n v riab7.e axon! Liiu cl)Tioue shock wave in
coma re^.::iOri rC..t i-O P1-1)0)
the airec.un of the axis.
Flow within a Con-
vergent, Conical Tube
r
i--.
C
The above sketch (Figure 3) shows the flow within a convergent,
conical tube up to the axis. As the distance from the axis becomes
smaller, the angle V of the oblique shock wave increases until, at the
axis, it becomes 90 degrees. The oblique shock wave at the edge has
thus become a plane shock wave at the axis. This flow pattern is, how-
ever, not stable and can only be produced in the wind tunnel with short
conical tubes. Besides, there is no regeneration of the full adiabatic
pressure. In the case of longer conical tubes, and in the case of the
slightest disturbance in the short conical tubes, the plane shock wave
moves upstream along the axis of the cone and stands in front of the
conical tube.
Figure 4
Plane Shock Wave
in front of Tube
.may
Figure 4 shows a plane shock wave in front of a tube. A variant of this
flow, the flow round the Busemanh ring, will not be discussed further
 Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
SECRET - US OFFICIALS ONLY
-7-
here, for it produces only a slightly better compression than the simple
plane shock crave.
A plane shock wave always occurs in the supersonic range when the velocity
is retarded.
Po Wl
The plane shock wave takes place within a plane perpendicular to the
direction of flow. In the shock plane the velocity decreases to subsonic
velocity, while the pressure and the temperature both increase.
 Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
Pk (working ON Pp (impact pressure)
pressure)
Air
'ressure
Tank
Impact Pressure
with Manometer
val '~
Nozzle Compression Shock
From enclosure 1 it Is apparent that the efficiency of compression in the
plane shock wave drops rapidly with an increase in the had : nuitb:r. The
plane shock wave can be utilized technically only when the :each number is
slightly over 1.0.
The i?:ultiple-Shock.Su.personic Diffusor
`'Oh; Sim11.e 'ltro-i_iae71si.ona7.. ulti. Ac:-;'.ock Supersonic Diffusor i.n I Icory
iN seevery of the multiple-shock supersonic diffusor:
E:: j The invention of the multiple-shock supersonic usor was
preceded by an observation by Stodola, which was correctly interpreted
by Prandtl who thus pointed out the way to the invention of the mu]_ciple-
shock supersonic diffusor. Stodola had measured impact pressures in a
Laval nozzle through which steam was flowing and. struck upon an
inexplicable irregularity. (Stodola: Gas Turbines)
Figure 7
Stodola's Exper-
imental Arrangem
Impact pressures measured along the axis of a Laval nozzle through which
steer was flowing showed. a regul r decrease in the direction of higher
hack numbers of flow. The regularity of this pressure drop was inter-
rupted Ly a sudden pressure rise close, boh:i.:ad the mouth of the Laval
nozzle. Prandtl interpreted this as proof of an oblique ::hock rave ..O-in.;
out from the mouth. The pitotube showed a higher pressure behind his
oblique shock wave. tJith the same Mach number of approach flow, there-
fore, a combination of an oblique shock wave (oblique shock moving from
the mouth of the nozzle) and a plane shock wave (in front of the impact
tube) produces a greater regeneration of pressure than a simple plane
shock wave alone.
 When in 1940 the question of a high pressure regeneration from a super-
sonic flow with greater Mach numbers became more and more pressing on the
Trommsdorff missiles (Lorin engines), and when the slight pressure
regeneration in the plane shock wave was such a handicap to development
that the entire project seemed to be jeopardized, Prandtl recalled the
observation of Stodola, and urged that the combination of an oblique
and a plane shock wave be investigated.
In pursuing these investigations, which were carried on simultaneously
in three different places, by Oswatitsch, Seiffarth,and Walchner in
Goettin en Busemann and his co-workers in Braunschweig, 0
the fact was established that such a combination of
two shocks actually does provide for considerably greater pressure
regeneration than could be attained with the single plane shock wave.
The oblique shock wave: If a supersonic flow is deflected by the angle,6
from its direction of flow by a cutting; edge, this unsteady deflec't'-.-xi
continues in the flour in the form of a shock wave at the angleV'.
Approved For Release 2007/08/10 : CIA-RDP80-00810A004600160009-8
 Approved For Release 2007/08/10 : CIA-RDP80-0081OA004600160009-8
The combination of several shocks in the multiple-shock diffusor: If the
flow still has supersonic velocity after the first oblique shock wave, it 25X1
can pass through a second oblique shock wave, and so on, until after a
final shock wave, which is mostly a plane shock wave, the flow finally
is decelerated to :ub:uonic velocity. In the question of what combinations
of shocks at a given Mach number and a given number of shocks produces
an optimal efficiency of supersonic compression, it can be derived and
^ 25X1
mathematical presentation; they can be determined only from
case to case through time consuming numerical and graphical operations.
The optimal compressions in these three-dimensional multiple-shock
diffusors are always a little better than the optimal compressions in the
corresponding two-dimensional multiple-shock diffusors, so that the
values of the two-dimensional diffusors give a very good indication of
the operational performance of all multiple-shock diffusors.
In Enclosure 2 the degrees of efficiency of optimal two-dimensional
multiple-shock diffusors, depending on the square of the Mach number of
the flow, are plotted, and in each case all the values for optimal
diffusors with two, three, and more shocks are combined into a single
curve. The degrees of efficiency for the two-dimensional shock wave,
already presented in Enclosure 1, are also plotted as a single shock
diffusor.
From an observation of the curves in Enclosure 2 it is apparent that
the degree of efficiency in the multiple-shock supersonic diffusor
at all Mach numbers can be brought close to the maximum value of
1.0, if the number of shocks is high enough.
The inverse Meyer-Prandtl flow as a diffusor with infinite shock
number: If the number of shocks is very high, whereby the value E
becomes smaller and smaller and finally differs only very little
from 1.0, then a very interesting flow pattern is produced, which
we shall call the "inverse Meyer-Prandtl flow". In his dissertation
(1905) Meyer investigated, upon Prandtl's suggestion, the constant
deflection of a supersonic flow with a'simultaneous transformation
of velocity into pressure. (A nozzle process in the sense of the
introductory section.),
(See Figure 9A gas flows along''a solid wall with supersonic velot-its. There is, on
the other side of the wall, ?lan.evacuated:space.. The wall ends in a cutting edge,
after which the gas flow is unconfined, The .individual gas particles now flow,
while. cooling arid expanding,. zrith increasing; velocity on curved paths into the
evacuatpd.space on the other side of the.solid walla
a. All streamlines are similar to one another only in respect to the
center of similarity at the point of the cutting edge S.
b. On a straight line (Mach line) drawn through the point of the
cutting edge (S) variables of state (pressure, density, tempera-
ture, velocity and Mach number) are always the same.
c. A straight line drawn through the point of the cutting edge S
bisects all other streamlines at the same angleo