ISO 17540:2025
(Main)Space systems — Liquid rocket engines and test stands — Vocabulary
Space systems — Liquid rocket engines and test stands — Vocabulary
This document provides terms and definitions for design, tests, reliability analysis and quality control of liquid rocket engines. The terms can be used in all types of documentation and subject-matter literature, related to standardization or use of the results of field-specific works.
Systèmes spatiaux — Moteurs de fusée liquides et stands d'essai — Vocabulaire
General Information
Relations
Standards Content (Sample)
International
Standard
ISO 17540
Second edition
Space systems — Liquid rocket
2025-10
engines and test stands —
Vocabulary
Systèmes spatiaux — Moteurs de fusée liquides et stands d'essai
— Vocabulaire
Reference number
© ISO 2025
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ii
Contents Page
Foreword .v
1 Scope . 1
2 Normative references . 1
3 Terms and definitions . 1
3.1 General .1
3.2 Major engine units .2
3.3 Engine types based on the cycle configuration.2
3.4 Engine types by reusability and multiple-firing capability .2
3.5 Engine types by purpose .3
3.6 Thrusters types based on the cycle configuration.3
3.7 Basic parameters and performance characteristics of the engine .4
3.8 Engine time characteristics, types of operating time and operating life .7
3.9 Thruster performance characteristics .8
3.10 Engine operating modes .9
3.11 Operating modes of liquid-fuelled thruster .10
3.12 Components of the combustion chamber and gas generator .10
3.13 Engine nozzle types . 12
3.14 Nozzle components . 12
3.15 Nozzle characteristics . 13
3.16 Nozzle operating modes .14
3.17 Nozzle flows .14
3.18 Turbopump: components .14
3.19 Pump characteristics . 15
3.20 Turbopump performance characteristics .16
3.21 Engine control devices .16
3.22 Devices and methods to create the engine steering efforts .17
3.23 Engine cooling .17
3.24 Engine thermal protection.18
3.25 Engine test: general . .18
3.26 Engine test types: thermal loads .18
3.27 Engine test types: by connection with the vehicle being moved .19
3.28 Engine test types: by test site .19
3.29 Engine test types: by organizational factor and test site .19
3.30 Engine test types: by test conditions .19
3.31 Engine test types: by accelerated data acquisition .19
3.32 Engine test types: by test purpose .19
3.33 Types of tests, specific for thrusters . 20
3.34 Engine test procedures .21
3.35 Engine test conditions .21
3.36 Test results . 22
3.37 Engine reliability . 22
3.38 Engine defects . 23
3.39 Engine failure modes . 23
3.40 Engine operation . 23
3.41 Engine health analysis .24
3.42 Engine reliability factors .24
3.43 Engine quality control .24
3.44 Structural and functional analysis of the engine reliability . 25
3.45 Test stands: general . 25
3.46 Test stand types . 26
3.47 Stand systems . 26
3.48 Post-test processing systems . . 28
3.49 Stand systems devices and components . 28
3.50 Stand room . 29
iii
Index .31
iv
Foreword
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This second edition cancels and replaces the first edition (ISO 17540:2016), which has been technically
revised.
The main changes are as follows:
— several terms are revised, including "rocket engine", "liquid rocket propulsion system", "high-altitude
firing test", etc.;
— redundant terms are deleted.
Any feedback or questions on this document should be directed to the user’s national standards body. A
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v
International Standard ISO 17540:2025(en)
Space systems — Liquid rocket engines and test stands —
Vocabulary
1 Scope
This document provides terms and definitions for design, tests, reliability analysis and quality control of
liquid rocket engines. The terms can be used in all types of documentation and subject-matter literature,
related to standardization or use of the results of field-specific works.
2 Normative references
There are no normative references in this document.
3 Terms and definitions
ISO and IEC maintain terminology databases for use in standardization at the following addresses:
— ISO Online browsing platform: available at https:// www .iso .org/ obp
— IEC Electropedia: available at https:// www .electropedia .org/
3.1 General
3.1.1
rocket engine
RE
reaction engine that contains within itself, or carry along with itself, all the substances necessary for its
operation or for the consumption or combustion of its fuel, not requiring any outside substance
3.1.2
liquid rocket engine
LRE
rocket engine (3.1.1) which uses liquid propellant
3.1.3
liquid-fuelled thruster
LFT
liquid-propellant propulsion device which is used as an actuator in launch vehicles, spacecraft and descent
vehicle control systems for acceleration, attitude control, stabilization, trajectory correction, rendezvous,
docking, braking, descent, landing and other manoeuvring operations, as well as for generating gravity by
imparting acceleration
3.1.4
liquid rocket propulsion system
LRPS
system consisting of LRE (3.1.2), propellant tanks, propellant (propellants) feeding system, thrust vector
control system and engine control system
3.1.5
clustered liquid rocket propulsion system
CLRPS
liquid rocket propulsion system (3.1.4) consisting of rocket engines of various applications, which are fed
from common propellant tanks, but have autonomous (independent) propellant feed systems
3.1.6
pressure-fed system
system in which propellants are supplied to the chamber (3.2.1) by their expulsion from propellant tanks
3.1.7
pump-fed system
system in which propellants are supplied from the tanks to the chamber (3.2.1) by pumps driven by a turbine,
which together form the turbopump unit
3.2 Major engine units
3.2.1
thrust chamber
chamber
engine unit where the liquid propellants are metered, injected, atomized, mixed, and burned to form hot,
gaseous reaction products, which in turn are accelerated and ejected at high velocity
3.2.2
turbopump
TPU
engine unit (component) designed to pump propellant into the chamber (3.2.1), gas generator (3.2.4) and
engine control devices and consisting of pumps (3.18.1) and the driving turbines (3.18.5)
3.2.3
booster turbopump
BTPU
auxiliary turbopump (3.2.2) of the engine designed to increase propellant pressure in the pipelines upstream
of the engine pumps (3.18.1)
3.2.4
gas generator
engine unit where propellant is chemically converted into exhaust gas
3.3 Engine types based on the cycle configuration
3.3.1
closed cycle engine
engine, in which exhaust gas of the gas generator, after having been used to drive the turbopump (3.2.2), is
fed to the chamber (3.2.1)
3.3.2
open-cycle engine
engine, in which exhaust gas of the gas generator, after having been used to drive the turbopump (3.2.2), is
released into the environment
3.4 Engine types by reusability and multiple-firing capability
3.4.1
expendable engine
engine, designed for a single use as intended
3.4.2
reusable engine
engine, designed for multiple use as intended
3.4.3
single-start engine
engine which performs a single firing during one use as intended
3.4.4
multiple-start engine
engine which performs multiple firings during one use as intended
3.5 Engine types by purpose
3.5.1
main engine
engine, designed to accelerate the vehicle being moved
3.5.2
correction engine
engine, designed to change the velocity when correcting the trajectory of the space vehicle being moved
during the coasting flight
3.5.3
control engine
engine, designed to control the velocity vector of the space vehicle being moved during the powered phase
3.5.4
retrorocket engine
engine, designed to reduce the velocity of the space vehicle being moved
3.6 Thrusters types based on the cycle configuration
3.6.1
catalytic liquid-fuelled thruster (LFT)
LFT (3.1.3) in which propellant is converted into gaseous chemical reaction products using a catalyst
3.6.2
thermo-catalytic liquid-fuelled thruster (LFT)
catalytic LFT (3.6.1) in which the catalyst is heated by the external heat source
3.6.3
electro-thermo-catalytic engine
thermo-catalytic LFT (3.6.2) which uses an electrical power source
3.6.4
radio-thermo-catalytic liquid-fuelled thruster (LFT)
thermo-catalytic LFT (3.6.2) which uses a radioactive power source
3.6.5
thermal liquid-fuelled thruster (LFT)
LFT (3.1.3) in which the conversion of propellant into gaseous combustion products and the increase in
exhaust velocity are achieved by heating the propellant with an external source of energy
Note 1 to entry: Power is supplied to the propellant or products of chemical reactions.
3.6.6
electro-thermal liquid-fuelled thruster (LFT)
thermal LFT (3.6.5) which uses an electrical power source
3.6.7
radio-thermal liquid-fuelled thruster (LFT)
thermal LFT (3.6.5) which uses a radioactive power source
3.6.8
electrolytic liquid-fuelled thruster (LFT)
monopropellant LFT (3.1.3) in which where the electrolysis of the propellant is part of the engine cycle
3.6.9
throttleable liquid-fuelled thruster (LFT)
LFT (3.1.3) which has a throttling device
3.7 Basic parameters and performance characteristics of the engine
3.7.1
rated engine operating conditions
set of rated engine parameter values according to the specification
3.7.2
propellant mass flow rate
ṁ
mass of a propellant passing through a corresponding feed line or valve per unit of time
3.7.3
oxidizer mass flow rate
ṁ
ox
mass of oxidizer passing through a corresponding feed line or valve per unit of time
3.7.4
fuel mass flow rate
ṁ
huel
mass of fuel passing through a corresponding feed line or valve per unit of time
3.7.5
propellant volume flow rate
Q
volume of a propellant passing through a corresponding feed line or valve per unit of time
3.7.6
oxidizer volume flow rate
Q
ox
volume of oxidizer passing through a corresponding feed line or valve per unit of time
3.7.7
fuel volume flow rate
Q
fuel
volume of fuel passing through a corresponding feed line or valve per unit of time
3.7.8
prestart flow rate
mass flow rate of a propellant during the period of time, starting from the moment the first firing command
is given until the thrust (3.7.10) builds up to the value, equal to 5 % of the nominal thrust
3.7.9
engine mass mixture ratio
r
m
ratio of the oxidizer mass flow rate (3.7.3) to the fuel mass flow rate (3.7.4) during engine operation
3.7.10
engine volume mixture ratio
r
v
ratio of the oxidizer volume flow rate (3.7.6) to the fuel volume flow rate (3.7.7) during engine operation
3.7.11
chamber pressure
p
c
average static pressure of combustion products at the entry section of the combustion chamber (3.12.1) near
the thrust chamber injector head (3.12.3)
3.7.12
gas generator pressure
p
gg
average static pressure of combustion products at the entry section of the combustion chamber (3.12.1) near
the gas generator injector head (3.12.4)
3.7.13
propellant combustion products exhaust velocity
exhaust velocity
c
e
exhaust flow velocity at the nozzle (3.12.12) exit plane of the engine, determined in a one-dimensional
approximation
3.7.14
engine reaction force
resultant of gas and hydrodynamic forces, acting on the internal surfaces of the engine during the expulsion
of combustion products
3.7.15
chamber thrust
F
result of the engine reaction force (3.7.14) and the ambient pressure forces acting on the engine external
surfaces, excluding external aerodynamic drag forces
3.7.16
thrust impulse
I
t
time integral of the chamber thrust (3.7.15)
3.7.17
engine afterburn impulse
engine thrust impulse (3.7.16) over a time interval that determines the engine shutdown (3.10.8)
3.7.18
engine specific impulse
I
sp
ratio of chamber thrust (3.7.15) to the propellant mass flow rate (3.7.2) of the engine (IR= /m )
s
Note 1 to entry: Thrust engine (chamber) specific impulse (3.7.13) is determined in vacuum (I ) and at sea level
sp,vac
(I ).
sp,sl
Note 2 to entry: engine specific impulse (3.7.18) also equals the derivative of the thrust impulse (3.7.16) with respect to
mass or volume of the usable propellant.
Note 3 to entry: For LFT the term “specific impulse” is used for steady-state continuous mode, single firings mode and
the steady-state impulse mode.
3.7.19
engine volume specific impulse
I
sp/V
ratio of the chamber thrust (3.7.15) to the propellant volume flow rate (3.7.5) of the engine (I = F/Q)
sp/V
Note 1 to entry: engine volume specific impulse (3.7.19) also equals the derivative of the engine thrust impulse (3.7.16)
with respect to the volume of usable propellant.
3.7.20
chamber thrust coefficient
CF
ratio of the chamber thrust (3.7.15) to the product of the stagnation pressure at the nozzle throat (3.14.8),
area and the nozzle flow coefficient (3.7.26)
3.7.21
chamber specific thrust coefficient
ƞ
i
ratio of the actual chamber specific impulse in vacuum to the ideal (design) value (3.7.25), defined by the
same values of mixture ratio, chamber pressure (3.7.11) and the nozzle geometric expansion ratio (3.15.1)
3.7.22
total chamber specific thrust coefficient
chamber specific thrust coefficient (3.7.21), defined at the mixture ratio, corresponding to the maximum ideal
(design) value
3.7.23
chamber flow characteristic
flow characteristic
product of the combustion pressure at a specified section of the thrust chamber (3.2.1) and the nozzle throat
(3.14.8), divided by the propellant mass flow rate (3.7.2) through the chamber
Note 1 to entry: the specified section of the thrust chamber (3.2.1) may be:
— the initial cross-section of the thrust chamber (3.12.1) near the injector head, when analysing stability of chamber
characteristics during serial production;
— the initial nozzle cross-section (3.14.7), when analysing multiphase flows (3.17.4).
3.7.24
thrust ratio
ratio of the chamber thrust to the product of the gas pressure at the given section of the thrust chamber
(3.2.1) and the nozzle throat (3.14.8) (KR=⋅/PA )
R ch ch thr
ch
Note 1 to entry: The thrust ratio also equals the ratio of the chamber-specificimpulse to the chamber flow characteristic
(3.7.23).
3.7.25
flow characteristic coefficient
ratio of the actual chamber flow rate characteristic (3.7.23) to the ideal (design) value, determined at the
same propellant mixture ratio and chamber pressure
3.7.26
nozzle flow rate coefficient
flow rate coefficient
ratio of the actual rate of gas flow through the nozzle (3.12.12) to the ideal (design) value, defined by the
same values of the temperature and stagnation pressure in the nozzle throat (3.14.8), gas constant and the
local adiabatic exponent
3.7.27
nozzle coefficient
ratio of the actual chamber thrust coefficient (3.7.20) in vacuum to the ideal (design) value, defined by the
same values of the mixture ratio, chamber pressure (3.7.11) and the nozzle geometric expansion ratio
(3.15.1) (ϕ =KK/ )
ch ts. s
3.7.28
chamber coefficient
ratio of the actual characteristic velocity (3.7.29) in the chamber to the ideal (design) value, defined by the
same values of the mixture ratio and the chamber pressure (3.7.11)
3.7.29
characteristic velocity
product of the stagnation pressure in the nozzle throat (3.14.8), the throat area, and the nozzle flow coefficient
(3.7.21), divided by the chamber mass flow rate (3.7.2)
3.7.30
chamber ideal parameter value
chamber parameter value, corresponding to the equilibrium flow of combustion products (gas generation
products) when there is no heat rejection and no friction
3.7.31
gas generator ideal parameter value
gas generator parameter value, corresponding to the equilibrium flow of combustion products (gas
generation products) when there is no heat rejection and no friction
3.7.32
wet mass
mass of the engine structure and propellants filling the pipelines and units
3.7.33
engine altitude characteristic
relation of the engine thrust (3.7.15) to the ambient pressure at constant values of the mixture ratio and the
chamber pressure (3.7.11)
3.7.34
engine throttle characteristic
relation of the engine thrust (3.7.15) to the chamber pressure (3.7.11) at constant values of the mixture ratio
and the ambient pressure
3.8 Engine time characteristics, types of operating time and operating life
3.8.1
propellant supply period
time interval from the moment of full opening of the solenoid valve (3.21.1) to the moment of full closure of
the solenoid valve
3.8.2
design life
period of time during which the engine is expected to operate within its specified design parameters
3.8.3
engine operating time
either the operating duration or the number of the engine cycles of operation, or both
3.8.4
engine service life
total engine operating time (3.8.3) accumulated during the design life (3.8.2), when the engine is used as
intended
3.8.5
engine service life limit
total engine operating time (3.8.3) upon reaching which the operation of the engine can be continued only
after a decision is made on the possibility to extend this limit
3.8.6
liquid-fuelled thruster (LFT) total service life limit
service life limit (3.8.5) with respect to firing duration for continuous and pulse modes
Note 1 to entry: In addition to the total service life limit for LFT (3.1.3) the service life limit is also determined by the
following:
— number of firings;
— impulse mode duration;
— continuous mode duration;
— total propellant consumption for catalytic LFT (3.6.1).
3.9 Thruster performance characteristics
3.9.1
liquid-fuelled thruster (LFT) total thrust impulse
LFT (3.1.3) thrust impulse, at which the mean integral value of thrust or chamber pressure (3.7.11) is greater
or equal to 0,9 of the steady-state value of the thrust or chamber pressure during one firing
3.9.2
liquid-fuelled thruster (LFT) partial thrust impulse
LFT (3.1.3)thrustimpulse at which the mean integral value of thrust or chamber pressure (3.7.11) is less than
0,9 of the steady-state value of the thrust or chamber pressure during one firing
3.9.3
liquid-fuelled thruster (LFT) unit impulse
LFT (3.1.3) thrust impulse per one firing when operating in the pulse or single firing mode
3.9.4
liquid-fuelled thruster (LFT) total impulse
LFT (3.1.3) thrust impulse during the service life limit
3.9.5
liquid-fuelled thruster (LFT) impulse
force action of LFT (3.1.3), characterized by time change of thrust or chamber pressure (3.7.11) during one firing
3.9.6
liquid-fuelled thruster (LFT) rated thrust
LFT (3.1.3) thrust in the steady-state continuous mode and under rated operating conditions
3.9.7
liquid-fuelled thruster (LFT) firing
time interval from the moment when voltage is supplied to the LFT (3.1.3) solenoid valve (3.21.1) until the
moment when voltage is turned off from the LFT
3.9.8
liquid-fuelled thruster (LFT) afterburn impulse duration
time interval from the moment when voltage is turned off from the LFT (3.1.3)electric valve until the
moment when thrust (3.7.10) or the chamber pressure (3.7.11) falls to the value equal to 0ю1 of the thrust or
the chamber pressure in the steady-state continuous operating mode
3.9.9
liquid-fuelled thruster (LFT) off-time
time interval from the moment when the voltage is removed from the LFT (3.1.3) solenoid valve (3.21.1) until
the moment when the voltage is supplied for the next firing
3.9.10
liquid-fuelled thruster (LFT) firing cycle
total of LFT firing time and LFT off-time (3.9.9)
3.9.11
liquid-fuelled thruster (LFT) firing frequency
value which is reciprocal of LFT firing cycle (3.9.10)
3.9.12
liquid-fuelled thruster (LFT) duty cycle
ratio of the LFT firing cycle (3.9.10) to the firing
3.9.13
barrier coefficient for liquid-fuelled thruster (LFT) pulse mode
ratio of the LFT firing (3.9.7) to the firing cycle
3.9.14
liquid-fuelled thruster (LFT) thrust build-up time
time interval from the moment when voltage is supplied to the LFT (3.1.3) solenoidvalve (3.21.1) until the
moment when thrust or chamber pressure (3.7.11) equals 0,9 of the value of the thrust or chamber pressure
in the steady-state continuous mode
3.9.15
liquid-fuelled thruster (LFT) thrust delay
time interval from the moment when voltage is supplied to the LFT solenoid valve (3.21.1) until the moment
thrust or chamber pressure (3.7.11) equals 0,1 of the value of the thrust or chamber pressure in the steady-
state continuous mode
3.9.16
liquid-fuelled thruster (LFT) propellant reaction delay
time interval from the moment when propellant is fed to the LFT chamber until the moment when the
chamber pressure (3.7.11) reaches the value, equal to the pressure when there is no propellant reaction
3.9.17
liquid-fuelled thruster (LFT) propellant ignition delay
time interval from the moment the second propellant enters the chamber until the ignition starts
3.9.18
liquid-fuelled thruster (LFT) propellant (fuel, oxidizer) average mass flow rate
ratio of the LFTpropellant mass flow rate (3.7.2), consumed per one firing, to the firing
3.10 Engine operating modes
3.10.1
engine operation
performance by the engine of operations necessary to generate thrust or change its values either in direction
or magnitude or both, and either provide operational conditions for the components of the moving vehicle or
ensure compliance with the established engine requirements or both
3.10.2
engine operating mode
set of the engine parameter values defined by the processes, occurring in the engine
3.10.3
engine main operating mode
engine operating mode (3.10.2) which is critical when performing the main task
3.10.4
engine start
engine operating mode (3.10.2) beginning from the moment the start command is issued and ending when
the main operating mode (3.10.3) is reached
3.10.5
steady-state operating mode
engine operating mode (3.10.2) when the mean thrust and mixture ratio values remain constant
3.10.6
transient operating mode
engine operating mode (3.10.2) in which the mean thrust and mixture ratio parameters vary over time
3.10.7
engine preliminary operating mode
steady-state operating mode (3.10.5) in which the thrust is lower than that of the main operating mode
Note 1 to entry: Engine preliminary operating mode is part of the engine start (3.10.4).
3.10.8
engine shutdown
engine operating mode (3.10.2) from the initial shutdown command to the cessation of thrust
3.10.9
engine final operating mode
engine steady-state operating mode (3.10.5) prior to the engine shutdown, when the thrust is less than the
thrust of the main operating mode
3.10.10
engine firing interval
time interval between the reusable engine (3.4.2) shutdown and the first command for the next firing
3.11 Operating modes of liquid-fuelled thruster
3.11.1
liquid-fuelled thruster (LFT) continuous operating mode
LFT (3.1.3) operating mode during one firing with the specific impulse value constant in time
3.11.2
liquid-fuelled thruster (LFT) pulsed operating mode
LFT (3.1.3) operating mode with multiple-firing (multiple firings) in which the specific impulse depends on
characteristics of each individual firing
Note 1 to entry: Minimum duration of the pulses is limited by the time taken for the thruster valves to open and close,
since this limits the repeatability of the process.
3.11.3
liquid-fuelled thruster (LFT) steady-state pulsed mode
LFT pulsed operating mode (3.11.2) when the pulse shape (3.9.5) becomes stable at the constant values of
firing frequency (3.9.11) and shutdown frequency
3.11.4
liquid-fuelled thruster (LFT) operating mode with connected pulses
LFT pulsed operating mode (3.11.2) in which thrust or chamber pressure (3.7.11) during pauses between
LFT firings (3.9.9) decreases to values greater than 0,1 of the steady-state thrust or chamber pressure in
continuous operation
3.11.5
liquid-fuelled thruster (LFT) cyclic operating mode
LFT (3.1.3) operating mode consisting of repeating combinations of continuous and pulsed modes or
repeating combinations of firings and off-time of different durations
3.11.6
liquid-fuelled thruster (LFT) single-firing mode
LFT (3.1.3) operating mode with off-time periods during which the engine goes back to its original
operating state
3.12 Components of the combustion chamber and gas generator
3.12.1
combustion chamber
part of the chamber (3.2.1) where propellants burn into combustion products
3.12.2
combustion chamber of a gas generator
part of the gas generator (3.2.4) designed for mixing and converting propellants into gas generation products
3.12.3
thrust chamber injector head
part of the thrust chamber (3.2.1) that is a device for feeding either propellants or gas generator exhaust
products, or both, into the combustion chamber (3.12.1) and for their initial mixing
3.12.4
gas generator injector head
part of the gas generator (3.2.4) that is a device for feeding propellants into the combustion chamber of a gas
generator (3.12.2) and for their initial mixing
3.12.5
bottom of the thrust chamber injector head
element of the thrust chamber injector head (3.12.3) that separates the cavities of the propellants or gas
generator products (propellants) from each other, or separates them from the combustion space and the
external environment
3.12.6
bottom of the gas generator injector head
element of the gas generator injector head (3.12.4) that separates the cavities of the propellants from each
other, or separates them from the combustion space and the external environment
3.12.7
injector
device for introducing propellants or gas generator products into either the combustion chamber or the gas
generator, or both
3.12.8
cooling tract of the chamber
set of channels in the chamber body and the thrust chamber injector head (3.12.3) providing either flow
cooling or transpiration cooling
3.12.9
cooling tract of the gas generator
set of channels in the chamber body and the gas generator injector head (3.12.4) providing either flow cooling
or transpiration cooling
3.12.10
chamber film cooling band
element of the thrust chamber (3.2.1) designed to introduce one of the propellants or gas generator products (one
of the propellants) into the near-wall region of the combustion zone to create a protective layer of liquid or gas
3.12.11
gas generator film cooling band
element of the gas generator (3.2.4) designed to introduce one of the propellants or gas generator products (one
of the propellants) into the near-wall region of the combustion zone to create a protective layer of liquid or gas
3.12.12
nozzle
part of the thrust chamber (3.2.1) that forms a variable cross-section channel, where the thermal energy of
combustion products is converted into kinetic energy of the exhaust jet.
Note 1 to entry: The engine nozzle may be either fixed or movable relative to the stationary parts of the thrust chamber
(3.2.1), and may also include a movable section for thrust vector control
3.13 Engine nozzle types
3.13.1
axisymmetric nozzle
nozzle (3.12.12) whose internal surface on the side of combustion products flow is symmetric relative to its axis
3.13.2
circular nozzle
axisymmetric nozzle (3.13.1) in which any cross-section of the combustion products flow perpendicular to
the symmetry axis is a full circle
3.13.3
conical nozzle
round nozzle (3.13.2), whose diverging part, beginning near the throat section, has a straight contour
3.13.4
contoured nozzle
nozzle (3.12.12) in which the expanding part has a curvilinear contour shaped to increase the efficiency of
the nozzle
3.13.5
annular nozzle
axisymmetric nozzle (3.13.1) in which some or all cross-sections of the combustion products flow,
perpendicular to the symmetry axis, are ring shaped
3.13.6
pin nozzle
annular nozzle (3.13.5), in which the contour of the expanding part almost entirely or completely lacks an
external section
3.13.7
dish-shaped nozzle
annular nozzle (3.13.5), in which the contour of the expanding part almost entirely or completely lacks an
internal section
3.13.8
extendable nozzle
nozzle (3.12.12) with one or several extendable skirts that, when deployed, serve as a continuation of the
expanding part of the nozzle
3.13.9
skewed nozzle
nozzle (3.12.12), exit of which is inclined to the nozzle axis at the angle, different from the right angle
Note 1 to entry: A skewed nozzle consists of the main axisymmetric part and a small non-axisymmetric part.
3.13.10
adjustable nozzle
nozzle (3.12.12) with a variable expansion ratio that can be changed during operation
3.14 Nozzle components
3.14.1
nozzle profile
line of intersection of the nozzle (3.12.12) surface with the plane, passing through the main axis
3.14.2
optimized nozzle profile
contoured nozzle profile, in which the expanding part is determined either through calculations or through
experiments, or both, to achieve the require engine performance characteristics
3.14.3
uniform characteristic nozzle profile
contoured nozzle profile, in which the expanding part provides a parallel flow at the nozzle exit section
(3.14.10) with a constant velocity at any point of the section
3.14.4
shortened nozzle profile
contoured nozzle profile, in which the expanding part corresponds to the initial section of the expanding
contour of a nozzle with a uniform characteristic
3.14.5
nozzle profile with an angular point
nozzle profile (3.14.1) with a break in the contour
3.14.6
nozzle exit section profile
closed line drawn through exit endpoints of all nozzle profiles (3.14.1)
3.14.7
initial nozzle cross-section
flow section of the chamber followed by a sharp reduction in the flow section area
3.14.8
nozzle throat
cross-section of the nozzle (3.12.12) which has minimum area
3.14.9
critical nozzle cross-section
flow section of the nozzle (3.12.12) where combustion products speed is equal to the local sound speed
3.14.10
nozzle exit section
engine nozzle cross-section, which is perpendicular to the central axis and which passes through endpoint
of the nozzle profile (3.14.1)
Note 1 to entry: For the annular nozzle (3.13.5), the exit section runs through the endpoints of the external area of the
nozzle profile (3.16.1); for the skewed nozzle (3.13.9) the exit section runs through the endpoint of the shortest contour.
3.14.11
nozzle converging section
part of the nozzle (3.12.12) between the initial nozzle cross-section (3.14.7) and the nozzle throat (3.14.8)
3.14.12
nozzle diverging section
part of the nozzle (3.12.12) between the nozzle throat (3.14.8) and the nozzle exit
3.15 Nozzle characteristics
3.15.1
nozzle geometric expansion ratio
ratio of the nozzle exit section (3.14.10) to the nozzle throat (3.14.8)
3.15.2
gas expansion ratio in the nozzle
ratio of the combustion product total pressure in the initial nozzle cross-section (3.14.7) to the static pressure
in the exit section
3.16 Nozzle operating modes
3.16.1
rated operating mode of the nozzle
operating mode of the nozzle, when the gas pressure at the exit section is equal to the ambient pressure
3.16.2
nozzle under-expansion mode
operating mode of the nozzle, when gas pressure in the exit section is higher than the ambient pressure
3.16.3
nozzle over-expansion mode
operating mode of the nozzle, when gas pressure in the exit section is lower than the ambient pressure
3.16.4
nozzle design altitude
altitude above the sea level where the nozzle operating mode is rated under standard atmospheric conditions
Note 1 to entry: Instead of the altitude above the sea level, it is allowed to use the corresponding ambient pressure
3.17 Nozzle flows
3.17.1
equilibrium nozzle flow
nozzle flow, characterized by power, chemical and phase balance of combustion products
3.17.2
non-equilibrium nozzle flow
nozzle flow, which lacks energy, chemical and phase equilibrium of combustion products or at least one of
these kinds of equilibrium
3.17.3
chemically frozen nozzle flow
nozzle flow, characterized by constant chemical composition of combustion products
3.17.4
multiphase nozzle flow
nozzle flow, characterized by the availability of gaseous, liquid and solid phases of combustion products
3.17.5
nozzle velocity lag
velocity difference of the condensed phase particle and the gaseous environment in the nozzle (3.12.12)
3.17.6
nozzle thermal lag
temperature difference of the condensed phase particle and the gaseous environment in a nozz
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