Nozzle Expansion Research Articles

NEAR FIELD FEATURES OF ORTHOGONALLY IMPINGING UNDER EXPANDED TRIANGULAR AND HEXAGONAL SUPERSONIC JETS

In this experimental investigation the shock structure and shear layer growth in the near field of non-circular supersonic jets impinging orthogonally on a flat surface were studied under two different nozzle expansion ratios (NPRs). The supersonic jets were of triangular and hexagonal cross-sectional shapes with equal area at the nozzle exit. The expansion ratio of all the nozzles used in the experiments was the same at 1.44, with a throat area of 75 mm2. This corresponded to a design exit Mach number of 1.8 for optimum expansion (NPR= 5.87). Experiments were conducted for three different impinging distances (LP) from the nozzle exit plane. The distances in the experiments were 2Dh, 4Dh, and 6Dh, where Dh was the hydraulic diameter of the nozzle exit shape. Experiments were carried out at NPR= 8 and 11(both under expansion conditions) at each impinging distance. Flow visualization of the impinging jet's shock structure was recorded using a Z-type Schlieren system employing concave mirrors which were polished to λ/6. The experiments were repeated for circular jets at the same experimental conditions and the shock structures of both circular and non-circular jets were compared. The shear layer thickness and its variation in the near field were obtained from the image processing of the schlieren images. The impinging distance was observed to have a significant effect on the growth of shear layer via modified shear layer shock interactions in the near field due to the presence of the orthogonal wall and its distance from the nozzle exit. The impinging plate was square in cross-section with a side dimension of 175 mm. The effect of impinging distance on the shock structure was found to be different for circular and non-circular jets. The jet shocks in the impinging jet, the plate shock in the impinging zone and the tail shocks near the high-speed radial wall jet exhibited a varied behavior depending upon the shape of the jet and its impinging distance. Hexagonal and triangular jets did not produce any Mach disk in the near field at shorter impingement distances such as Lp=2Dh. The introduction of the impinging effect and subsequent reduction in Lp decreased the Mach disk distance from the nozzle exit of both shapes, till the Mach disk disappeared at Lp=2Dh. The length of the jet shocks near the nozzle exit also decreased with an increase in Lp for both circular and non-circular shapes. A decrease in Lp impeded the shear layer growth along the jet boundary in the near field for all jet shapes. The triangular jet exhibited a thicker layer along the jet boundary originating from the corner, when compared to the other two jet shapes.

Advancing the development of supersonic natural gas liquefaction through understanding heavy hydrocarbon crystallization mechanism

During supersonic natural gas liquefaction, the solidification of heavy hydrocarbon molecules can be used to improve the efficiency of the liquefaction. This study investigates the supersonic crystallization of n-hexane using an experimental system designed by ourself for precise measurement of liquid heavy hydrocarbon vaporization. The results demonstrate that n-hexane gas condenses rapidly 1.5 cm downstream of the nozzle throat, releasing significant latent heat. The subsequent crystallization of these droplets transitions the substance fully from liquid to solid, and the latent heat released during crystallization is much lower than that during condensation. Under different conditions, we found that higher partial pressures intensify n-hexane crystallization, accelerating liquid phase growth despite slower solid phase mass increase. Furthermore, larger nozzle expansion ratios shorten the time and space distribution of the liquid phase mass fraction. Molecular dynamics simulations reveal n-hexane’s melting and surface crystallization temperatures to be approximately 160 ± 1 K and 158 ± 1 K, respectively, with a supercooling degree of 2 K, explaining the observed crystallization behavior in supercooled droplets.

CFD investigation of supersonic two parallel expansions nozzles

This paper aims to validate the design contours and result parameters of the newly developed Dual Expansion Nozzle (DEN) using the Fastran software when it enters the non-adaptation regime with the increase in altitude and variation of the Nozzle Pressure Ratio (NPR). DEN design and Computational Fluid Dynamic (CFD) application utilize the High Temperature (HT) model defined by a calorically imperfect and thermally perfect gas, providing good accuracy compared to PG model. The computational domain is decomposed into subdivisions of structured grids using the algebraic grid generator software CFD-GEOM. The mesh convergence was analyzed using three mesh resolutions: coarse, medium, and fine. A comparison was made between the two conventional nozzles, MLN and the newly developed BPN, as both are currently used in aerospace propulsion to enhance aerodynamic performance. If a rocket engine operates under strongly over-expanded conditions with ambient pressure considerably higher than the nozzle exit pressure, and by increasing NPR value, the flow separates from the wall and can lead to high side loads. These loads are reduced in DEN. The flow separation point is observed close to the exit section for the DEN in comparison with Minimum Length Nozzle (MLN) and Best Performance Nozzle (BPN) for the same NPR, resulting in low flow separation and low side load effects for DEN. This result demonstrates a small loss in the exit Mach number, leading to less loss in the thrust coefficient and a small effect of boundary layer friction for DEN. Enhanced over-expanded aerodynamics is observed in DEN compared to MLN and BPN. The application is made for air with an exit Mach number equal to 3.00.

Performance analysis of the adaptive cycle engine with cooling air: From propulsion performance, thermodynamic efficiency, exergy and infrared characteristics

The adaptive cycle engine gains significant attention due to its variable thermodynamic cycles and multi-duct characteristics, which are utilized for thermal management to enhance engine performance. This paper explores the features of multi-duct technology in adaptive cycle engines and aims to improve engine flight performance. An analysis is conducted on the impact of duct air bleed on engine propulsion characteristics, thermodynamic cycle properties, and exhaust system infrared radiation characteristics. Initially, a component-level model of the adaptive cycle engine was established, considering CDFS duct, bypass duct, and third duct air cooling devices. This model uses the ducted cold air to cool the temperature of low-pressure turbine exit, center cone wall, and main nozzle expansion section wall. Subsequently, a method for analyzing the infrared radiation of the engine exhaust system was proposed based on the above model, and the performance of engine components and the overall engine was analyzed in terms of energy, exergy, mechanical efficiency, and thermal efficiency. Finally, the numerical simulations of the propulsion performance, thermodynamic efficiency and infrared characteristics of the adaptive cycle engine with cooling air were carried out under conditions of ground takeoff, subsonic cruise, and supersonic cruise. The simulation results show that the cooling air can effectively enhance engine performance and infrared stealth capabilities under various flight conditions. The ducted air extraction measures proposed in the paper could effectively improve the propulsion efficiency and stealth characteristics of the engine, providing important references for the development of more efficient aeroengine.

Selection of propulsion characteristics for systematic assessment of future European RLV-options

AbstractThe propulsion system as the key-element of any space transportation concept is systematically investigated supporting a study on potential future European RLV. Different liquid propellant combinations are compared and evaluated. Subsequently, main stage rocket engines are defined for the two cycle options: open gas-generator and closed staged-combustion. Four different propellant combinations are considered all based on liquid oxygen (LOX) as oxidizer and with the fuel options liquid hydrogen (LH2), liquid methane (LCH4), liquid propane (LC3H8) and kerosene (RP1). These combinations result in eight different generic engines with sub-variants using different nozzle expansion ratios in first and upper stage application. Engine characteristics are similar, as far as conceivable, to be well-suited for the system level comparison of pre-defined RLV-launchers. However, characteristics are not necessarily identical as different engine architectures and propellants might make individual choices necessary. Engine performance characteristics are compared with similar existing engines when available. It is shown that closed-cycle staged combustion engines bring significant performance gains, particularly in sea-level operations. On the other hand, hydrocarbon- and open-cycle gas-generator engines offer a better thrust-to-weight-ratio than hydrogen and staged combustion cycle.

Study on the key parameters of ice particle air jet ejector structure

Existing ice particle jet surface treatment technology is prone to ice particle adhesion during application, significantly affecting surface treatment efficiency. Based on the basic structure of the jet pump, the ice particle air jet surface treatment technology is proposed for the instant preparation and utilization of ice particles, solving the problem of ice particle adhesion and clogging. To achieve efficient utilization of ice particles and high-speed jetting, an integrated jet structure for ice particle ejection and acceleration was developed. The influence of the working nozzle position (Ld), expansion ratio (n), and acceleration nozzle diameter ratio (Dn) length-to-diameter ratio (Ln) on the ice particle ejection and acceleration was systematically studied. The structural parameters of the ejector were determined using the impact kinetic energy of ice particles as the comprehensive evaluation index, and the surface treatment test was conducted to verify the results. The study shows that under 2 MPa air pressure, the ejector nozzle parameters of n = 1.5, Dn = 4.0, Ld = 4, and Ln = 0 mm can effectively eject and accelerate the ice particles. The aluminum alloy plate depainting test obtained a larger paint removal radius and resulted in a smoother aluminum alloy plate surface, reducing the surface roughness from 3.194 ± 0.489 μm to 1.156 ± 0.136 μm. The immediate preparation and utilization of ice particles solved the problems of adhesion and storage in the engineering application of ice particle air jet technology, providing a feasible technical method in the field of material surface treatment.

Optimization of exhaust ejector with lobed nozzle for marine gas turbine

To attain high-performance ejector configurations, an ejection characteristic testing system was established initially to validate the reliability of the Realizable k-ε turbulent model. Subsequently, optimization investigations were conducted on lobed nozzle ejectors with various structural parameters. The effects of four key structural parameters, including lobed nozzle expansion angle α, lobed nozzle width d, number of lobes in the nozzle n, and height of the square-to-circle section h, were systematically studied. Furthermore, the CRITIC method was employed for multi-objective evaluation to identify the optimal design configuration for the casing ejector. The research findings revealed that among the structural parameters, the lobed nozzle expansion angle α exerted the greatest influence on the ejection coefficient and pressure loss coefficient. The weights of the evaluation criteria were determined by the CRITIC method as follows: ejection coefficient (49.38%) < pressure loss coefficient (50.62%). The optimal design configuration determined by the CRITIC method included α = 45°, d = 150 mm, n = 14, and h = 600 mm. The resulting enclosure design ensures smooth airflow within the system, preventing the backflow of high-temperature mainstream fluid and heating the enclosure. It also maintains a temperature distribution in the typical cross-section that meets specified requirements. Additionally, it facilitates improved mixing of mainstream and secondary fluid and reduces exhaust gas temperature.

Two parallel expansions for improving supersonic axisymmetric nozzle performance

The aim of this work is to develop a numerical computation program allowing designing new contours of a supersonic axisymmetric nozzle having two expansions at the throat, named by DEN (Dual Expansion Nozzle). This new nozzle gives a uniform and parallel flow at the exit section, to improve considerably the performances compared to the conventional Minimum Length Nozzle (MLN), and the Best Performances Nozzle (BPN). The present nozzle has a two unknowns external and central body curved walls. Each of them is started by an initial expansion angle to give a uniform and horizontal flow at the exit section. Two others transition regions are calculated in parallel with the contours points to give the desired exit Mach number. The walls are determined point by point by the High Temperature Method of Characteristics (HT MOC) model. The resolution of the four compatibility and characteristics equations is done numerically by the finite difference predictor corrector algorithm. The validation of the results is controlled by the convergence of the calculated critical sections ratio to that given by the theory. The design depends on four parameters, where MLN and BPN become special cases of DEN. A comparison is made with MLN, since it is currently used in the aerospace propulsion and with BPN aiming to improve their performances. The comparison is made for the same critical mass flow rate. The results demonstrate a remarkable reduction up of 45 %, and 52 % in the mass of DEN when the exit Mach number ME = 3.00 and the stagnation temperature T0 = 2000 K. The application is made for air and for future aerospace missiles in order to improve their trajectory parameters. The chosen example demonstrates an improvement of 13 % and 16 % on the missile range compared, respectively to MLN, and BPN.

Numerical modeling of the non-equilibrium condensation in multi-component flows through the high expansion ratio nozzles

Excessive flow expansion in the thruster nozzles reduces the flow temperature and pressure. This provides conditions for the condensation phenomena in the nozzle, which can affect the thrust performance of thrusters. In this paper, the condensation effect of water components in the decomposition products of hydrogen peroxide in a typical thruster nozzle is modeled. To describe the state and properties of the real gas mixture, the Peng-Robinson equation of state is used. Moreover, the effects of H2O2 concentration and nozzle profile on the condensation and nozzle performance parameters are studied. The results show that by more flow expansion and crossing the saturated state of condensable components, the thrust coefficient improves up to 6%. This improvement decreases by increasing the H2O2 concentration in the reaction chamber for a fixed nozzle expansion ratio. Due to the small growth of droplets during the free molecular regime of the droplet growth in the high-speed flow, the non-equilibrium condensation continues until the nozzle outflow. In addition, although all the profiles have the same expansion ratio for nucleation start, increasing the nozzle length results in larger nano-droplets, more liquid mass fraction formation up to 4%, and moving the Wilson point to the further upstream sections in the nozzle.

Numerical study on dynamic flow of flexible extension nozzle in rocket motors

The use of extension nozzles in rocket propulsion systems is one of the effective methods to increase the geometric expansion ratio and thrust of the nozzle. As a new type of extension nozzle, the flexible extension nozzle can achieve continuous changes in nozzle expansion ratio. To explore the characteristics of the flow field structure at the “bulge” structure and the step structure in the actual configuration of the flexible extension nozzle, and the impact on the performance of the nozzle, a three-dimensional numerical simulation model of the flexible extension nozzle was first established to study the flow field at the “bulge” structure, and then a two-dimensional unsteady numerical simulation model was established by using the coupling technology of dynamic grid and overlapping grid to study the flow field at the step structure. The results show that during the unfolding of the nozzle surface, a complex flow structure of “expansion fan-mixing layer-shock-vortex” were formed in the step area at the junction of the fixed section and the moving extension section. The specific impulse loss of the 3D model is larger than that of the 2D model due to the consideration of the “bulge” structure. The flow loss of the “bulge” structure and the step structure to the nozzle is 0.73% and 0.65% respectively.

Design optimization of a low-cost three-stage launch vehicle with modular hybrid rocket motors

This article investigates the impact of modular propulsion system design on the performance and cost of a three-stage hybrid rocket. Furthermore, it conducts a multi-objective optimization of unit payload cost, take-off mass, and payload mass ratio, considering factors such as the number of motors and layout considerations. The optimization design scheme for the three-stage hybrid rocket is divided into four cases. In the first case, each stage is equipped with a fixed single motor, and each stage is independently optimized without modular design. The second case considers the use of multiple motors in the first and second stages, still without modular design. The third case also involves multiple motors in the first and second stages, but all motors in each stage have identical parameters except for the nozzle expansion ratio, implementing a modular design. In the fourth case, the number and layout of the motor design method are the same as those in the third case, with independent optimization in the third stage using partial modular design. The results indicate that the unit payload cost of the multi-motor non-modular design case can be reduced by 13.12% compared to the single-motor non-modular design case. Within the modular case, the full modular design case is slightly inferior to the partial modular design case. Based on the above data, it can be concluded that the first and second stages of modular rockets offer the best performance and the lowest cost.

Nitrous Oxide/Ethanol Cylindrical Rotating Detonation Engine for Sounding Rocket Space Flight

There are few experimental studies on rotating detonation engines (RDEs) with liquid propellants. This study reveals the static thrust performance of a cylindrical RDE with ethanol and liquid nitrous oxide as propellants under atmospheric pressure. This RDE had an inner diameter of 40 mm, a maximum combustor length of 230 mm, a nozzle contraction ratio of 1.7, and a nozzle expansion ratio of 9.1. Nineteen experiments were conducted at total mass flow rates of 184–210 g/s, mixture ratios of 3.6–5.9, and combustion pressures of 0.35–0.46 MPa, resulting in a maximum detonation velocity of 1840 m/s (approximately 80% of the theoretical detonation velocity, 2251 m/s), maximum thrust at sea level of 294 N, and maximum specific impulse at sea level of 148 s. In addition, the maximum characteristic exhaust velocity, c*, was 1716 m/s, which was 99% of the theoretical value. The characteristic length of the combustion chamber at this time was 0.15 m. Since conventional rocket combustion requires 1.57 m to achieve the same c* efficiency, this study shows that detonation combustion can reduce the combustor size by 88%.

Analysis of internal pressure and vibration of a flexible expansion ratio elongation nozzle shape

The nozzle made of flexible materials can expand and elongate its profile at different flight altitudes, greatly improving the engine’s endurance and ballistic accuracy. However, this also introduces the problem of structural vibration in the nozzle, which adversely affects engine performance. This paper introduces a high-precision hyperelastic material constitutive model for the flexible elongation nozzle and uses numerical simulation to conduct a detailed analysis of the internal pressure and structural vibration under aerodynamic pressure loads. As the flight altitude increases, the pressure magnitude and distribution on the inner wall of the same elongated flexible expansion section do not change. The pressure distribution on the inner wall of the flexible elongation expansion section exhibits certain nonlinear characteristics, and its magnitude decreases gradually along the axial position. The vibration of the elongated expansion nozzle profile becomes more intense with the increase in flight altitude. At high characteristic flight altitudes, the elongation length of the longer expansion section approaches an undamped vibration state.

Commissioning of Upgrades to T6 to Study Giant Planet Entry

The scientific potential of a mission to the ice giants is well recognized and has been identified by NASA and ESA as a high priority on several occasions, most recently in the 2023–2032 Decadal Survey. The payload capacity of such a spacecraft is limited by the heat shield thickness, which must be sized conservatively due to a lack of reliable data for convective and radiative heat flux along the proposed entry trajectories. Major upgrades to the Oxford T6 Stalker Tunnel have been commissioned that allow study of giant planet entry trajectories, including a flammable gas handling system, a Mach 10 expansion nozzle, and a steel shock tube with optical access. Initial testing has been completed in shock tube and expansion tunnel modes, with peak shock speeds of 18.9 km/s achieved. Convective heat flux and surface pressure were measured at several locations on a 45° sphere cone model in expansion tunnel mode. Measurements of the radiating shock layer were made in shock tube mode to assess the effect of CH4 concentration. This work establishes the first high-enthalpy giant planet entry test bed in Europe.

Large eddy simulation for jet noise characteristics of liquid rocket engine with guide groove

The study of jet noise characteristics of liquid rocket engines is a prerequisite for evaluating the noise level of rockets and effectively controlling the noise. To obtain the effect of engine operating and structural parameters on jet noise, the Large Eddy Simulation combined with the FW-H integration method is applied for the scaled engine with guide groove under different chamber pressures and nozzle expansion ratios. The simulation results show that the noise field distribution of the jet with a guide groove is highly symmetrical and has obvious directivity. The noise caused by jet impingement on the guide groove is strongly correlated to the distribution of high vorticity regions. Remarkably, higher chamber pressure leads to longer potential jet length, higher noise sound pressure levels, and greater enhancement effect of impact on noise. However, the frequency spectral characteristics are basically invariant. When the chamber pressure is constant, the increment of the nozzle expansion ratio in a certain range results in lower noise peak frequency and weaker impact effect.

Experimental study on the physical mechanism of pulse noise in the tail of a recoilless gun

During the firing process of a recoilless gun, there is a high-intensity pulse noise phenomenon in its rear breech area. In order to obtain the physical mechanism of the pulse noise formation, this article designed experiments and used high-speed cameras and overpressure sensors to discover a “three peaks” phenomenon in the pulse noise at the shooter’s position. Using resistance strain sensors, the changes of strain with time in the chamber and nozzle expansion section of the recoilless gun during the firing process were recorded. By comparing the micro-strain data and overpressure data, the shock waves at the shooter’s position can be categorized into two types: internal shock waves and external shock waves. Internally, the opening of the blocking sheet initiates the formation of the initial shock wave, followed by the formation of a second pressure wave within the combustion chamber. The peak-to-peak interval between these two shock waves is nearly fixed, and the ratio of the dual peaks increases with the distance of the measurement position. The third shock wave forms externally to the nozzle, induced by the explosion of unburned propellant. As the weapon’s distance from the ground increases, the peak values of the overpressure data decrease, and the waveform of this explosion-induced shock wave transitions from a single-peak phenomenon to a double-peak phenomenon. This study can provide a reference for the study of the potential risk of injury to operators using single soldier recoilless guns and similar weapons.

Velocity measurements in particle-laden high-enthalpy flow using non-intrusive techniques

This research aims at analysing the particle-laden flow of the hypersonic high-enthalpy wind tunnel L2K, situated in Köln at the German Aerospace Center (DLR). In the L2K wind tunnel, Martian atmosphere can be created, and the facility can simulate heat load conditions encountered during atmospheric entry of Martian missions. In the tests, a simplified Martian atmosphere (97% CO2 and 3% N2) was used. The high-enthalpy flow was loaded with micrometric particles of magnesium oxide. The particles’ mean velocity was measured with a 2D–2C particle image velocimetry (PIV) system, in the region right downstream the nozzle expansion of the wind tunnel. The work proves the possibility of creating a high-enthalpy particle-laden flow for thermal protection systems (TPS) testing with simulated Martian atmosphere. Average particle velocities of around 2000 m/s are measured and compared with the numerical simulation of the wind tunnel’s particle-free flow, and with the flow velocity measured with tunable diode laser absorption spectroscopy (TDLAS). The study also highlights some unexpected results and features of the high-enthalpy particle-laden flow and proposes some theories for the causes of such effects, which include agglomeration due to melting, and gravitational effect.

Dynamic Modelling and Simulation of Industrial Plant Fire and Gas Detection Systems; Parametric Analysis of Nozzle Flow

Industrial plant safety is of paramount importance, and the effective operation of Fire and Gas Detection Systems (FGDS) plays a critical role in preventing and mitigating incidents. This study presents a comprehensive parametric computational fluid dynamics (CFD) analysis of nozzle flow, aimed at enhancing our understanding of the complex fluid dynamics within nozzle geometries. Nozzles find wide applications in industries ranging from aerospace propulsion systems to industrial processes. The analysis employs CFD simulations with varying geometric parameters, such as nozzle shape, throat diameter, and expansion ratio, to investigate their influence on flow characteristics, including velocity profiles, pressure distributions, and shock formations. The parametric study involves systematic variations of these geometric parameters, allowing for the identification of optimal configurations for specific applications. The results shed light on the intricate interplay between nozzle geometry and flow behavior, providing valuable insights for nozzle design and optimization.

The influence of gas heating on the calculated supersonic parameters of gas-powder flow at slag blower in the converter. Message 1

The article shows that one of the priority tasks in steel production technology is to increase the durability of the lining of oxygen converters. It is shown that the slag blower is a radical way to increase the durability of the lining of oxygen converters. When blowing the melt with oxygen, the MgO content in the slag is 6-8%. To increase the chemical affinity of the slag and the lining, it is proposed to blow the converter slag with jets of nitrogen or powder gas. To achieve optimal chemical affinity of the lining and slag, it is recommended to modify the latter by increasing the content of magnesium oxide in it to 12-14%. Increased durability of the lining makes it possible to solve the problem of replacing wasteful wastes of coal slag and changing environmental concerns on the excess media. It is shown that an attempt to increase the efficiency of modeling very complex processes of interaction of supersonic jets with converter slag has not yet been solved a. A critical analysis of existing methods for constructing mathematical models based on the use of known gas-dynamic laws of interaction of free turbulent jets with the melt is carried out. It has been established that only supersonic jets with their characteristic shock-wave structure always enter the slag melt. An integral method for calculating the parameters of a gas-powder flow in a supersonic nozzle taking into account the powder concentration is presented. It is determined how heating of the gas suspension affects the required pressure in front of the nozzle block, the speed and flow density in the outlet section of the supersonic nozzle. The gas dynamics system is connected to the heat exchange of the flowing gas-dispersed mixture in the Laval expansion nozzles to establish the heating effect in front of the nozzle block of the two-phase gas suspension to a temperature that will ensure The increase in the fluidity of the gas flow is doubled and its kinetic energy is increased threefold. The method of expanding the parameters of the gas-powder flow between the Laval expansion nozzle involves the infusion of 10 physical effects into 14 parameters of the gas-powder flow. It is shown that numerical expansions allow the tuyere body to be used as a heat exchanger for heating the gas suspension, as well as for increasing the pressure of the gas-powder jet that flows from the nozzles

A novel relaxation drift model for simulating liquid-vapor momentum non-equilibrium in two-phase ejectors

A novel relaxation drift model (RDM) was proposed to simulate the interphase momentum non-equilibrium (slip) inside two-phase work-recovery CO2 ejectors. The RDM was validated using the experimental data of 5 different ejectors under 15 trans-critical conditions in total. The proposed model was compared with the existing homogeneous model (HM) and algebraic slip model (ASM) through case studies. The effects of ejector mixer design and nozzle expansion state were also disclosed. The results showed that the prediction errors of ejector performance by the RDM were within 17 %, which was smaller than the maximum error of 23 % predicted by the HM and the ASM. The numerical results demonstrated that the liquid droplets had lower velocities than the vapor when the liquid-vapor mixture accelerated, but the droplets travelled faster than the vapor during deceleration of the mixture. It was also revealed that the velocity slip would affect the shock pattern at the diffuser inlet if the mixer was choked. The effects of slip, however, became insignificant when the mixer was designed too large. In addition, the RDM might overestimate the ejector performance if the motive nozzle was strongly under-expanded. This study facilitates understanding of interphase velocity slip in two-phase CO2 ejector flow.