ISO 17540:2016

INTERNATIONAL ISO
STANDARD 17540
First edition
2016-12-15
Space systems — Liquid rocket
engines and test stands — Terms and
definitions
Systèmes spatiaux — Moteurs de fusée liquides et stands d’essai —
Termes et définitions
Reference number
ISO 17540:2016(E)
©
ISO 2016

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ISO 17540:2016(E)

COPYRIGHT PROTECTED DOCUMENT
© ISO 2016, Published in Switzerland
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ISO 17540:2016(E)

Contents Page
Foreword .v
Introduction .vi
1 Scope . 1
2 Terms and definitions . 1
2.1 General . 1
2.2 Engine units . 1
2.3 Engine types by way of work process . 2
2.4 Engine types by multiplicity of use and integration . 2
2.5 Engine types by purpose . 2
2.6 Low-thrust engine types by way of work process . 3
2.7 General parameters and performance of engine . 3
2.8 Engine time characteristics, types of operating and resources . 6
2.9 Low-thrust engine performance . 7
2.10 Engine operation modes . 9
2.11 Low-thrust engine operation modes .10
2.12 Chamber (gas generator) components .10
2.13 Gas generator types .12
2.14 Operating process in chamber (gas generator) .12
2.15 Nozzle types .14
2.16 Nozzle items .15
2.17 Nozzle characteristics .16
2.18 Nozzle operation modes .16
2.19 Flow in nozzle .16
2.20 Turbine pump components . .17
2.21 Pump characteristics .18
2.22 Turbine pump general characteristics .18
2.23 Automation units .19
2.24 Devices and methods of control efforts creation in engines .19
2.25 Engine cooling .19
2.26 Engine thermal protection .20
2.27 Engine tests: General .20
2.28 Types of engine tests: Thermal loads .21
2.29 Types of engine tests: Associate with rocket .21
2.30 Types of engine tests: Test site .21
2.31 Types of engine tests: Organizational factor and test site .21
2.32 Types of engine tests: Test conditions .21
2.33 Types of engine tests: Accelerated data accessing .22
2.34 Types of engine tests: Test purposes .22
2.35 Types of tests specific for low-thrust engines .23
2.36 Test technology .23
2.37 Test conditions .23
2.38 Test results.24
2.39 Engine reliability .24
2.40 Engine defects .25
2.41 Engine failure modes .25
2.42 Engine operation .25
2.43 Analysis of engine technical status .26
2.44 Engine reliability index .26
2.45 Engine quality control .26
2.46 Structural and functional analysis of reliability .27
2.47 Test stands: General .27
2.48 Stand types .28
2.49 Stand systems .28
2.50 Post-test processing.29
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ISO 17540:2016(E)

2.51 Stand system elements.30
2.52 Stand compartments .31
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ISO 17540:2016(E)

Foreword
ISO (the International Organization for Standardization) is a worldwide federation of national standards
bodies (ISO member bodies). The work of preparing International Standards is normally carried out
through ISO technical committees. Each member body interested in a subject for which a technical
committee has been established has the right to be represented on that committee. International
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ISO collaborates closely with the International Electrotechnical Commission (IEC) on all matters of
electrotechnical standardization.
The procedures used to develop this document and those intended for its further maintenance are
described in the ISO/IEC Directives, Part 1. In particular the different approval criteria needed for the
different types of ISO documents should be noted. This document was drafted in accordance with the
editorial rules of the ISO/IEC Directives, Part 2 (see www.iso.org/directives).
Attention is drawn to the possibility that some of the elements of this document may be the subject of
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Technical Barriers to Trade (TBT) see the following URL: www.iso.org/iso/foreword.html.
The committee responsible for this document is ISO/TC 20, Aircraft and space vehicles, Subcommittee
SC 14, Space systems and operations.
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ISO 17540:2016(E)

Introduction
This International Standard is intended to be applied for all types of rocket engines which use a liquid
propellant.
The terms in this International Standard are specified in scope of design, testing, reliability analysis
and quality control of liquid rocket engines.
The terms are intended to be required for use in all types of documentation and literature including in
scope of standardization or using results of this activity.
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INTERNATIONAL STANDARD ISO 17540:2016(E)
Space systems — Liquid rocket engines and test stands —
Terms and definitions
1 Scope
The International Standard provides terms and definitions in scope of design, testing, reliability
analysis and quality control of liquid rocket engines. The terms are required for use in all types of
documentation and literature including in the scope of standardization or using the results of this
activity.
2 Terms and definitions
2.1 General
2.1.1
rocket engine
RE
reaction engine producing thrust for vehicle movement with the help of substances and energy sources
contained within the vehicle being moved
2.1.2
liquid rocket engine
LRE
rocket engine (2.1.1) using propellants in liquid form
2.1.3
low-thrust engine
LTE
rocket engine (2.1.1) of a thrust not more than 5 000 N
2.1.4
liquid rocket propulsion system
propulsion system including engine, propellant tanks, avionics for control sub-systems, pressure
vessels and control devices for pneumatic and hydraulic control sub-systems, propellant feed system,
actuators for steering sub-systems, and auxiliary equipment
2.1.5
clustered engine
liquid rocket propulsion system (2.1.4) consisting of multiple rocket engines (2.1.1), common propellant
tanks, and autonomous (independent) propellant feed systems
2.2 Engine units
2.2.1
chamber
engine assembly where propellant and/or gas generation products, as a result of chemical reactions,
are converted into products of combustion, created at the expiration of the reactive force
2.2.2
turbo-pump
TP
engine component designed to pump propellant into the chamber (2.2.1), gas generator sets and
automatic engine
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ISO 17540:2016(E)

2.2.3
booster turbo-pump
BTP
turbo-pump (2.2.2) engine support designed to increase propellant pressure in the pipelines to pump
(2.20.1)
2.2.4
gas generator
unit of engine wherein propellant, as a result of chemical reaction, is converted in gaseous products of
reaction at relatively low temperature
2.2.5
automatic engine controller
engine assembly designed for automatic control, regulation or maintenance of engine
2.3 Engine types by way of work process
2.3.1
engine with afterburning
engine where gas generation products after their use are used to drive the turbo-pump (2.2.2) assembly
2.3.2
engine without afterburning
engine where gas generation products after their use to drive the turbo-pump (2.2.2) assembly are
released into the environment
Note 1 to entry: Engine without afterburning have a pump (2.20.1) or a pressurized fuel supply.
2.3.3
single-mode engine
engine with one major mode
2.3.4
multimode engine
engine with several basic modes
2.4 Engine types by multiplicity of use and integration
2.4.1
expendable engine
engine intended for a specific purpose and used only one time
2.4.2
nonexpendable engine
engine intended for a specific purpose and used multiple times
2.4.3
single-start engine
engine started only once for a specific purpose
2.4.4
multi-start engine
restartable engine
engine started multiple times for one specific purpose
2.5 Engine types by purpose
2.5.1
main engine
engine intended to accelerate the space vehicle
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ISO 17540:2016(E)

2.5.2
correction engine
engine intended to correct the speed during the correction of trajectory of the space vehicle
2.5.3
control engine
engine intended to control the correction of the vector of the space vehicle in the active phase of the
trajectory of motion
2.5.4
retrorocket engine
engine intended to reduce the speed of the space vehicle
2.6 Low-thrust engine types by way of work process
2.6.1
catalytic engine
LTE (2.1.3) where the transformation of propellant into gaseous chemical reaction products is
performed with the help of a catalyst
2.6.2
thermo-catalytic engine
catalytic LTE where the catalyst is heated by the external heat source
2.6.3
electro-thermo-catalytic engine
thermo-catalytic LTE using an electrical source of energy
2.6.4
radio-thermo-catalytic engine
thermo-catalytic LTE using a radioactive source of energy
2.6.5
thermal engine
LTE (2.1.3) where the conversion of propellant in the gaseous products of chemical reactions is affected
by heating the fuel from an external source of energy which increases their rate of expiration
Note 1 to entry: Energy is fed to the propellant or products of chemical reactions.
2.6.6
electro-thermal engine
thermal LTE using an electrical energy source
2.6.7
radio-thermal engine
thermal LTE using a radioactive energy source
2.6.8
electrolytic engine
one-component of the LTE (2.1.3) where the electrolysis of the propellant is part of operating process
2.6.9
adjustable engine
low-thrust engine (2.1.3) that has a device to change the thrust
2.7 General parameters and performance of engine
2.7.1
rated performance
set of nominal values of the engine designated in the specifications
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2.7.2
mass flow rate
mass of fluid passing a specified line or gate in unit time
2.7.3
volume flow rate
volume of fluid passing a specified line or gate in unit time
2.7.4
pre-start consumption
propellant mass consumption during the time interval from the first start command until the thrust
build-up to a specified value equal to 5 % of the nominal
2.7.5
mixture ratio
ratio of oxidizer mass flow rate (2.7.2) to the fuel mass flow rate
2.7.6
volume ratio
ratio of oxidizer volume flow rate (2.7.3) to the fuel volume rate
2.7.7
pressure
average static pressure of combustion products at the beginning of the combustion
chamber (2.12.1) at the mixing system chamber
2.7.8
pressure
average static pressure of gas generation at the beginning of the combustion chamber
(2.12.2) at the mixing system gas generator
2.7.9
combustion temperature
stagnation temperature of combustion products at the exit from the combustion chamber
(2.12.1)
2.7.10
combustion temperature
stagnation temperature of gas generation at the exit from the gas generator (2.2.4)
2.7.11
exhaust velocity
velocity of exhaust stream through the nozzle (2.12.16) or a reaction engine, relative to the nozzle
2.7.12
engine reactive force
gas and fluid flow resultant force acting on the thrust chamber internal surfaces resulting from the
combustion gases
2.7.13
engine thrust
resultant of the engine reactive force (2.7.12) and the environment pressure forces acting on the engine
external surfaces (excluding external aerodynamic drag forces)
2.7.14
engine impulse
time integral of engine thrust
2.7.15
cut-off impulse
impulse (2.9.5) of engine thrust for the time interval defining the engine tail-off
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ISO 17540:2016(E)

2.7.16
specific impulse
 
R
ratio of engine thrust to the mass flow of propellant I =
 
s

m
 
Note 1 to entry: Thrust engine (chamber) specific impulse is converted in a vacuum and at sea level.
Note 2 to entry: Thrust engine (chamber) specific impulse is also an equalled derivative from the thrust engine
(chamber) impulse by weight or volume of propellant consumed.
Note 3 to entry: For LTE (2.1.3), the term “specific impulse” is used for steady-state continuous mode, single
inclusions mode and the steady-state impulse mode.
2.7.17
volume specific impulse
 
R
ratio of engine thrust to the propellant volume flow rate I =
 
sv,

v
 
2.7.18
thrust coefficient
ratio of chamber thrust to the product of the nozzle stagnation pressure (or chamber total pressure at
nozzle inlet) and the area of nozzle throat
2.7.19
coefficient of specific impulse
ratio of actual specific impulse to the theoretical value that is defined by the same values of mixture
ratio (2.7.5), the nozzle stagnation pressure or chamber total pressure at nozzle inlet
2.7.20
total coefficient of specific impulse
coefficient of specific impulse (2.7.16) defined at the mixture ratio (2.7.5) to be the maximum ideal value
2.7.21
consumable complex of chamber
consumable complex
product of the combustion pressure in a given section of the chamber (2.2.1) to a nozzle throat area,
referred to the mass flow of the propellant in chamber
Note 1 to entry: Given section of the chamber (2.2.1) is in analysis of camera characteristics stability during
serial production [initial section of combustion chamber (2.12.1) at (near) mixing system (2.12.3)] and in analysis
of multiphase flows (2.19.4) [initial section of nozzle (2.12.16)].
2.7.22
thrust complex
ratio of engine thrust chamber pressure and the product of combustion products in a given section of
the chamber (2.2.1) for an area of minimum section of the nozzle (2.12.16)
Note 1 to entry: Thrust complex is also equal to the ratio of camera-specific impulse to consumable complex
(2.7.19).
2.7.23
coefficient of consumable complex
ratio of the actual spending of the complex chamber rocket engine to the ideal that defined the same
values of the ratio components fuel pressure in the chamber (2.2.1)
2.7.24
coefficient of nozzle flow
coefficient of flow
ratio of the actual flow of gas through the rocket engine nozzle to the theoretical value, as defined under
the same temperature and total pressure in the nozzle throat, under the conditions for the gas constant
and the local adiabatic exponent
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2.7.25
nozzle coefficient
ratio of the actual thrust coefficient in a vacuum to the ideal that defined the same values of the mixture
ratio (2.7.5) and combustion pressure in the chamber (2.2.1) and the geometric expansion ratio nozzle
2.7.26
chamber coefficient
ratio of the real characteristic velocity in the chamber (2.2.1) to the ideal defined by the same values of
the mixture ratio (2.7.5) and the combustion chamber pressure
2.7.27
characteristic velocity
product of the nozzle stagnation pressure and nozzle throat area, referred to the mass consumption of
propellant in chamber
2.7.28
ideal parameter value
parameter value of chamber (2.2.1), corresponding to the equilibrium flow of combustion
products in the absence outlet heat and friction
2.7.29
ideal parameter value
parameter value of gas generator (2.2.4), corresponding to the equilibrium flow of
products gas generation in the absence outlet heat and friction
2.7.30
wet mass
mass of engine designed with propellants and other consumption articles filling its pipelines and
aggregates
2.7.31
relative mass
ratio of the wet mass (2.7.30) to the maximum thrust on the main steady-state operation
2.7.32
engine altitude characteristic
dependence of the thrust rocket engine on the environment pressure at constant values of the ratio of
the propellant components and the pressure in the chamber (2.2.1)
2.7.33
engine throttle characteristic
dependence of the engine thrust from the chamber pressure at constant values of the mixture ratio
(2.7.5) of propellants and the ambient pressure
2.8 Engine time characteristics, types of operating and resources
2.8.1
period of propellant flow
time interval from the moment of complete opening of the solenoid valve until it is completely closed
2.8.2
designed operating life
period of time during which the engine is expected to operate within its specified design parameters
2.8.3
engine operating time
operation duration and/or operation cycle number of the engine
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ISO 17540:2016(E)

2.8.4
engine verification time
mean time engine specified in the request for the proposal
2.8.5
engine specified resource
engine operating time (2.8.3) specified in the request for the proposal
2.8.6
engine working resource
total running time of engine during a specified period of service, used as directed
2.8.7
engine single working resource
work resource of engines, or part thereof, during one cycle operation
2.8.8
engine designated resource
total operating time after the expiry of which the use of the engine should be stopped
2.8.9
LTE total designated resource
operation duration assigned for continuous and pulse modes
Note 1 to entry: In addition to total designated resourse, for LTE (2.1.3), it is also determined designated resource
according to the following:
— number of inclusions (2.9.8);
— duration at impulse mode;
— duration at continuous mode;
— total propellant consumption for catalytic LTE.
2.9 Low-thrust engine performance
2.9.1
full thruster impulse
thruster impulse of LTE (2.1.3) at which the mean integrated value of thrust, or chamber pressure, is
more or equal to 0,9 of the steady-state value of the thrust, or chamber pressure, for the firing
2.9.2
part-thrust impulse
thruster impulse of LTE (2.1.3) at which the average integral value of thrust, or pressure (2.7.7) in the
chamber (2.2.1), is less than 0,9 the steady-thrust, or pressure in the chamber, at a switch
2.9.3
unit impulse
thruster impulse of LTE (2.1.3) or one firing (on-time (2.9.10)) in the pulse or single firing operation mode
2.9.4
total impulse
thruster impulse of LTE (2.1.3) over the operating duration
2.9.5
impulse
forceful impact of LTE (2.1.3) characterized by changes in traction or pressure (2.7.7) in the chamber
(2.2.1) at the time of a switch
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2.9.6
rated thrust
designed thrust level in a steady-state condition mode under nominal working conditions
2.9.7
conditional rated thrust
rated thrust of LTE (2.1.3) in a vacuum at an initial temperature of 288 K where structures and the
geometric expansion ratio of the nozzle is equal to 50
2.9.8
inclusion
on-time
time interval from the moment of voltage being applied to the thruster electric valve up to the moment
of reenergizing the LTE (2.1.3)
2.9.9
aftereffect
thruster electric valve reenergizing up to the moment when the thrust of the chamber pressure fall to a
value equal to 0,1 of the thrust of the chamber pressure in the steady-state continuous operation mode
2.9.10
off-time
pause between inclusions
time interval from the moment of the thruster electric valve reenergizing up to the moment of the next
voltage being applied
2.9.11
cycle period
on-time (2.9.8) and off-time (2.9.10) sum
2.9.12
inclusion frequency
reciprocal of cycle period (2.9.11)
2.9.13
cycle period to on-time ratio
duty cycle
reciprocal of duty cycle
2.9.14
coefficient of fill cycle operation
inclusion relation of LTE (2.1.3) to switching cycles
2.9.15
thrust build-up time
time interval from the ignition signal to the moment when the thrust or chamber pressure reaches a
value of 90 % of the steady-state thrust or the chamber pressure
2.9.16
thrust delay
time interval from the cut-off signal until the thrust or chamber pressure decreases to 10 % of steady-
state thrust or chamber pressure
2.9.17
propellant expansion delay
interval time from the start entry of the second component of propellant cell LTE (2.1.3) until the
pressure (2.7.7) in the chamber (2.2.1) reaches a value equal to the pressure in the absence of fuel
decomposition
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