Jet Exhaust Research Articles

Unsteady characteristics of supersonic base flow influenced by the exhaust jet

The supersonic base flow with an exhaust jet is significantly unsteady, whose characteristics can be influenced by the exhaust jet states. By the exhaust jet, we mean the working fluid produced by the rocket engine and exhausted by the base nozzle in supersonic velocity, which is used to generate thrust. A numerical investigation of the base flow at Mach 1.96 is carried out via the delayed detached eddy simulation. The effects of presence, expanded states, and temperature of the exhaust jet on the unsteady characteristics are analyzed. Compared with the exhaust jet off, the exhaust jet presence causes the base pressure fluctuation intensity to increase in the high-frequency band (SrD > 0.5) and changes the dominant mode from wake oscillation (SrD = 0.12) to the periodically expand-shrink of the exhaust jet diameter (SrD = 0.16). Compared with the underexpanded state, the overexpanded exhaust jet induces the flow separation inside the nozzle, which causes a decrease in the base pressure fluctuation intensity in the high-frequency band (SrD > 1.0), and the separation point movement and exhaust jet oscillation constitute the dominant mode (SrD = 0.13). Compared with the low-temperature jet, the base pressure fluctuation intensity is slightly reduced in the high-frequency band (SrD > 1.0) for the high-temperature jet, and the flapping motion of the jet shear layer becomes the dominant mode (SrD = 1.5).

Infrared radiation characteristics simulation of exhaust suppression type ships

An equivalent calculation method for the infrared-suppressed ship’s infrared radiation is carried out, which couples thermal characteristics of the ship surface and the exhaust system. Firstly, a refined physical model for the infrared suppressed exhaust system is proposed, which considers the processes of convective heat transfer and radiative heat transfer. The model generates the equivalent heat flow boundary conditions of the internal heat source through the numerical calculation of Computational Fluid Dynamics (CFD), and combines with the energy balance physical model. The energy balance model on the ship surface is developed, which considers external thermal environment conditions including itself radiation, solar radiation, atmospheric radiation, and reflective radiation among surface elements. The temperature calculation model for the ship with a jet exhaust system is formed. Subsequently, an exhaust-suppressed ship infrared radiation calculation model is formed based on the infrared radiation rendering of solid targets and the narrow spectral principle of exhaust plumes, which considers the spontaneous radiation and environmental radiation of ships. Finally, the infrared simulation results are compared and verified based on experimental data. The results indicate that the infrared radiation error on the ship’s typical parts is less than 30 %. Additionally, 90.1 % of the cabin walls using exhaust suppression systems have a temperature difference of less than 3 K under different operating conditions. It suggests that the exhaust suppression system is effective in reducing the temperature of the cabin wall. The calculation method for ship infrared characteristics constructed in this paper has reasonable results, and it can be used for infrared characteristic analysis of exhaust suppression ships.

Design, assembly and testing of an upgraded aeroacoustics wind tunnel facility

The Aeroacoustics Laboratory in the Department of Aerospace Engineering at Penn State currently serves several areas of experimental research. Most notable in the laboratory is the Anechoic Chamber of which the Aeroacoustics Wind Tunnel is an integral part. In recent years the dominant areas of research have included model aircraft exhaust jets and model rotors for helicopter applications involving detailed noise experiments. In the jet noise area, the applications have included aircraft in both flyover and take-off configurations. In the flyover application the air flow surrounding the exhaust jet has a significant effect on the radiated noise and must be simulated to produce experimental results that will be most useful in preliminary design of noise suppression applications. The Aeroacoustics Wind Tunnel at Penn State serves the purpose for this simulation. During aircraft take-off, the aircraft airspeed is in a range exceeding 250 ft per second, or Mach number of over M = 0.22. In the years preceding 2022, the velocity of the free jet flow of the Penn State facility was limited to less than 200 ft/sec. This paper reports on a project to design upgrades to the existing wind tunnel to improve the test section flow velocity to values closer to the 250 ft/sec target. The approach began with a preliminary design of the return flow ducting to convert the open jet, open return wind tunnel to a closed return tunnel. The concept is to make use of the flow energy in the exhaust of the tunnel to boost the input energy to the inlet fan that drives the flow to the test section. An integral part of the preliminary design involved making use of an upgraded flow analysis computer code to predict the gain in test section velocity of the new facility. For the available horsepower of two fans in the facility, the configurations of the existing and upgraded wind tunnel were entered into this analysis code based on a quasi-one-dimensional formulation. The code included empirical data formulated to include the various shapes of the components of the tunnels. The relative flow velocities in each section used simple incompressible continuity to calculate all velocities in the component sections. The average dynamic pressure in each section is calculated from the flow (constant) density and the velocity squared. Each section had a non-dimensional pressure loss parameter calculated from empirical data (from the relevant literature). The basic code (using the Excel algorithms) was based on a predictor-corrector concept in which the calculation was initialized with the estimated velocity and pressure of the flow at the entrance to the test section. The loss in each following section is estimated to determine the loss in the total pressure and the cumulated loss compared to the gains made at the two fans in the circuit. The initial guess is adjusted based on the total pressure value at the end of the circuit. During the related experiments, measurements of the total and dynamic pressure were performed at several joints between sections and compared to the values predicted in the analysis code. Various settings of the horsepower of the drive fans were used and the results were compared to flow measurements for the tunnel configurations (predominantly for the open return and closed return tunnel setups available before and after the upgrade project.) Results of the analysis and experiments are presented and analyzed together with the remaining activities planned for the coming months.

Jet Noise and Wing Installation Effects of Circular, Beveled and Rectangular Nozzles

AbstractWith the growth of modern turbofan engines, their integration under the wing becomes challenging and induces aerodynamic and acoustic interactions between the jet exhaust and the airframe. Jet noise reduction techniques have been widely studied over the past decades but their efficiency has still to be demonstrated once installed. The present lab-scale jet experiments at Mach 0.6 compare the noise radiated by beveled and rectangular installed nozzles to circular ones on a quarter-sphere radiation map using a microphone antenna. For all radiation angles, modified nozzles show an amplitude decrease of the jet-plate interaction tones of the noise spectra attributed to a strong coupling between the jet shear layers and the sound scattering at the plate trailing edge. Beveled nozzles achieve a noise reduction for all radiation angles with a maximum decrease up to 2 dB at receiver locations perpendicular to the plate. While rectangular nozzles show a similar behavior, a sound increase is observed for listeners parallel to the plate when the height-to-width ratio is small.

Modeling the air quality impact of aircraft emissions: is area or volume the appropriate source characterization in AERMOD?

Modeling dispersion of aircraft emissions is challenging because aircraft are mobile sources with varying emissions rates at different elevations depending on the operating mode. Aircraft emissions during landing and take-off cycle (LTO) influence air quality in and around the airport, and depending on the number of aircraft operations and location of the airport, this influence may be significant. AERMOD (v22112) incorporates a variety of conventional source types to characterize the intended emissions source, leaving the question of which conventional source type(s) best characterizes aircraft activities across the four modes of LTO cycle, unanswered. Currently, the publicly released version of FAA’s Aviation Environmental Design Tool (version 3e) models aircraft emissions as a set of AREA sources for all flight segments. A research version of AEDT allows users to model aircraft sources—both fixed wing and rotorcraft—as a series of VOLUME sources in AERMOD. However, both source treatments do not account for plume rise of aircraft jet exhaust. This paper compares AERMOD’s performance in describing SO2 concentrations associated with airport sources by comparing model results from the two source options during the summer campaign of the Air Quality Source Apportionment study conducted at the Los Angeles International Airport. We conclude that both VOLUME source and AREA treatments overestimate the highest observed SO2 concentrations despite not accounting for background sources. The VOLUME source option reduces this overestimation by using a higher initial plume spread than the AREA option does, and through the inclusion of meander. Our results suggest the need to include the plume rise of jet exhaust when using AERMOD for airport air quality studies.

Numerical study of jet-plate interaction and its effects on aeroacoustics

In modern under-wing-mounted high-bypass engines, the exhausted jet from the engine will interact strongly with the wing and high-lift devices. In this paper, simulations of an isothermal jet flow at Mach number Ma = 0.5 and close to a flat plate are carried out using large-eddy simulation. Far-field acoustic results are obtained via the surface integral method of Ffowcs Williams and Hawkings (FW-H). The studied cases include an isolated jet and other three configurations with the jet and a plate spaced at the distance H/D j = 1.0, 1.5, and 2.0, respectively, from the jet axis. Validation against experimental data is carried out. Far-field acoustic results show that the overall noise levels are significantly increased for the installed cases, particularly for polar angles in upstream of the nozzle. The noise levels are increased more as the plate approaches the jet axis, indicating the installation effect is related to the H/D and the polar angle. The pressure fluctuations are shown to decay along the radial direction. The mean flow and power spectral density of the axial velocity show the characteristics of the jet flow is not significantly changed by the presence of the plate.

Quantitative Color Schlieren for an H2–O2 Exhaust Jet Developing in Air

Throughout many decades, the Schlieren visualization method has been mainly used as means to visualize transparent flows in a qualitative manner. The images recorded provide data regarding the existence of the flow, or illustrate predicted flow geometries and details. The colored Schlieren method has been developed in the late 1890s and has always had the intent to provide quantitative data rather than qualitative pictures of the studied phenomena. This paper centers on applying a quantitative color Schlieren method to help determine the gasodynamic parameters of an H2–O2 exhaust jet, developing in air. A comparison between the parameters obtained through calibrating the color filter for the Schlieren method and the results from a CFD simulation is performed to assess the range of the CS (color Schlieren) measurement. This paper’s findings address the issues of calibrated color filter Schlieren encounter during its implementation and discusses possible errors appearing when the method is applied to a 3D flow. While the qualitative Schlieren images are still impressive to observe, the quantitative Schlieren presents challenges and a low measurement accuracy (75%) when applied to 3D flows and compared to 2D cases found in the literature (97–98%).

Evaluation of feasibility of commercial supersonic flight based on aeroacoustics

The drive to attain ever higher speeds, to be able to travel ever faster fuels the research and development for a commercial supersonic aircraft. This has previously led to the Concorde which travelled at more than twice the speed of sound. Now, in addition to business considerations about economic viability, supersonic aircraft must be quieter and emit less emissions. Considering the 20 years that have elapsed since Concordes retirement, this study aims to reevaluate the current challenges and limitations to achieving commercial supersonic flight again, in the context of noise. Identifying sonic booms and jet exhaust noise as two main challenges, it reviews current shape optimization methods, plasma as a sonic boom mitigator, sonic boom circumvention, chevron nozzles, variable cycle engines, engine positioning, and their corresponding limitations. Some of the methods have been refined for use and application in final stage design and manufacture of certain supersonic aircraft which indicates a certain feasibility.

Experimental Investigation of a Micro Turbojet Engine Chevrons Nozzle by Means of the Schlieren Technique

In connection with subsonic jet noise production, especially regarding the hot jet from a micro turbojet engine, we encountered a lack of recent high-resolution data in the literature describing the flow field using experimental validation through optical diagnoses. The objective of this paper is to examine and compare the influence on shear layers of the exhaust plug nozzle of a micro turbojet engine with and without chevrons mounted, using a high-speed camera used in Schlieren-type optical system diagnosis. Three different operating regimes are examined for both the baseline configuration and the configuration with 16 triangular-shaped chevrons. In conjunction with the image captures, the sound pressure level was recorded with the help of a microphone placed perpendicular to the flow, 0.4 m from the exhaust of the nozzle which was further processed. In quantitative terms, we found that the OASPL decreases by more than 1% when the engine is operating at higher regimes. Moreover, we found that the average exhaust jet angle, which is a measure of the quality of the fluid mixing layer is increased by 5% with respect to the baseline nozzle. By using the “darkest pixel” technique in Schlieren imaging, we can verify experimentally, for all working regimes, the theory that asserts that subsonic jet noise is a consequence of fine-scale homogeneous turbulence. Additionally, the potential novelty lies in the specific observations related to consistent dispersion of fine-scale eddies and how the presence of chevrons amplifies this uniformity within the turbulent field.

Designing jet deflector configuration for a semi-cryogenic rocket engine

This study deals with the design of a suitable configuration of jet deflector for supersonic exhaust gases from a semi-cryogenic launch vehicle engine using computational fluid dynamics. Computational model of combustion in semi-cryogenic engine, with kerosene as fuel and liquid oxygen as oxidizer, was developed in which exhaust gases from engine were impinged on the deflector surface. Different design configurations of deflector structure, obtained by the combination of various impingement angles of 15°, 20°, 25°, 30°, and 45°, and different exit radii of 5000, 10 000, and 20 000 mm, were used in the developed computational models to analyze impingement, deflection, and associated flow properties of exhaust gases, such as acoustics, force on deflector imparted by jet, temperature, and Mach number of flow field. A new ablation model was developed based on Vieille's law to determine the amount of refractory material ablated from deflector structure for different configurations. The developed models of ablation, combustion, and flow impingement were also validated with existing literature. It was found that the configuration with 20° impingement angle and 10 000 mm exit radius had ablation, acoustics, force, and flow properties in desired limit. Furthermore, to find the optimum uplift angle for designing jet deflector, configurations with uplift angles of 0°, 5°, 15°, and 25° were studied using the computational models developed in the study. It was observed that the configuration with 20° impingement angle, 10 000 mm exit radius, and 15° uplift angle was best suited for impingement and deflection of exhaust jet from the specified semi-cryogenic engine.

Lab-scale research on acoustically transparent stacks for hot exhaust jets in cooler cross-flow

From the exhaust stack of open cycle gas turbine power stations, the sound radiated is strongly refracted in the near-field by the thermal and velocity gradients of the jet. Research undertaken by the authors has identified that the presence of a cooler cross-flow leads to a bent-over plume downwind of the exhaust stack and a significant increase in sound in the far field by up to 10dB downwind when compared to spherical spreading. This paper is the first in a series that explores a unique “acoustically transparent silencer,” which prevents this by separating the sound from the flow. The concept of the approach is demonstrated in this paper, with aero-acoustic simulations conducted on a supercomputer and small-scale experiments in a wind tunnel. The numerical simulations have predicted the effects of using an acoustically transparent exhaust stack to reduce the refraction of sound in the near field, which reduced the lobes in the acoustic directivity downstream of the exhaust by up to 4.5 dB. This was further supported by experimental results in a wind tunnel with a reduction of up to 4 dB. These results were used to assist in the second iteration of designs for the Refrastack in a subsequent paper.

Modeling and Dynamic Stability Analysis of Distributed Electric Propulsion Tilt-Rotor UAV

Distributed electric propulsion tilt-rotor unmanned aerial vehicles (TRUAVs) with a lift-increasing flap in the propulsion wake have better aerodynamic efficiency than traditional tilt-rotor unmanned aerial vehicles and better takeoff/landing capacity than fixed-wing aircraft. However, for this configuration, establishing the propulsion–aerodynamic coupling model between the flap and the duct is difficult. This paper presents a flight dynamics model with propulsion–aerodynamic coupling components and analyzes dynamic characteristics during the transition period for TRUAVs with a distributed-propulsion/flap configuration. The development of the flight dynamics model considers the descriptions of the duct model, the lift-increasing flap model, and the propeller model. The model of the lift-increasing flap in both the duct exhaust jet and the freestream is the key component of the vehicle dynamics equations and is validated with experimental data and computational fluid dynamics results. Based on the dynamics model, the transition corridor of the tilt-rotor aircraft with multiple asynchronous tilting components and a geometrical constraint is calculated and analyzed. Additionally, the longitudinal dynamics stability of the aircraft in different flight speeds and configurations is assessed based on the numerical differentiation method. The results showed that the vehicle flight speed has more influence on longitudinal short-period stability than the angle of attack (AOA) of the flap. Also, the longitudinal long-period stability is affected by both the vehicle flight speed and the AOA of the flap. Finally, a mathematical derivation is developed to analyze the aircraft’s unstable state. The presented analysis results show that the stability of the long period is affected by the velocity derivative of the fuselage’s lift.

Experimental Investigation on Aluminum-Based Water Ramjet for Propelling High-Speed Underwater Vehicles

Major challenges in developing and realizing a novel aluminum–water reaction-based water ramjet propulsion system for high-speed underwater vehicles and demonstration of a water-breathing jet propulsion test facility are investigated. Two stages of combustion, propellant grain combustion and subsequent water combustion, with primary combustion products are adopted. High-pressure-molded propellant grains up to 45% of micro–nano () aluminum were prepared and combusted in the primary chamber, which exhibits mild ignition delay, and a residue of 4–6% was retained. Once water is injected into the secondary chamber, the net thrust generation is increased more than twice from the exhaust jet and improves the specific impulse by 40%. The lean fuel conditions in the secondary chamber lead to reduction in combustion propensity, which causes drop in efficiency. The ultrafine iron-oxide-catalyzed micro–nano blended propellants marginally improved the propulsive performance than the uncatalyzed compositions. The efficiency of the catalyzed propellants was enhanced up to 38.6%. Aluminum agglomeration in primary combustion considerably occurred; apparently, only a fraction of aluminum particles or agglomerates are completely burnt within the secondary chamber, and the remaining aluminum particles are either partially burnt or go unreacted.

Material selection based on joule heating simulation for resistojet thruster

Resistojets generate heat by passing an electric current through a resistive element, thereby heating a propellant at high pressure, which is subsequently expanded through a nozzle resulting in a supersonic exhaust jet. Specific impulse of a resistojet decides the performance of the thruster. Hence the goal here is to achieve highest specific impulse. The specific impulse of the resistojet is limited by the thermal design of the heater and the material selection. Since the design specifications on the resistojet model has been confirmed with calculated data points obtained from literature work, the material selection is done with the heat transfer simulation. Also, the materials selected as the simulation model is done on accordance with highest melting point and the factor of feasibility is taken into account. For achieving the heat transfer distributions among the flow channels through conduction allows the calculation of heat flux along the radial direction. The transient state condition of the turbulent fluid region is solved under the CFX solver. The simulation is done on different materials varying their thermophysical properties. The end result will be decided on the basis of the heat transfer rate and the residual obtained during higher iterations. Thus, this paper focuses on the choice of materials and the thermal analysis of the same to achieve the highest specific impulse and an efficient performance output.

Experimental Analysis on the Operating Line of Two Gas Turbine Engines by Testing with Different Exhaust Nozzle Geometries

This paper presents a special analysis study about the gas turbine operating line, and an overall description of a gas turbines project, based on experimental data from two particular applications, in order to convert two different types of aero engines into the same engine configuration. The experimental works were carried out with the aim of converting an Ivchenko AI-20K turboprop and a Rolls-Royce Viper 632-41 turbojet into free turbine turboshaft engines, to be used in marine propulsion, and also to obtain an experimental database to be used in other gas turbine applications. In order to carry out the experimental work, the engines were tested in turbojet configuration, to simulate the free turbine load by using jet nozzles with different geometries of the outlet cross-section. Following the engines’ tests, a series of measured data were obtained, through which it was possible to experimentally determine the operating line of some engine components such as the compressor, turbine, and exhaust jet nozzle. This paper is comprehensive and useful through its scientific and technical guidelines, the operation curves coming in handy in the thermodynamic analysis and testing methodology for researchers dealing with similar applications.

Steady-state sound propagation through hot exhaust jets in cooler cross-flow: A computational study

An analysis has been carried out to investigate the sound radiation through a heated jet in cooler cross-flow, which is representative of many industrial exhaust systems, using a hybrid steady-state computational fluid dynamics and computational acoustic model. The mean flow and temperature fields are modelled using steady-state computational fluid dynamics, with the turbulence modelled using Reynolds Averaged Navier-Stokes equations. The corresponding mean flow and temperature fields are used in the computational sound propagation model using linearised acoustic wave equation with mean flow based on a scalar flow potential. The results obtained from the computational simulations show that the flow significantly changes the sound propagation path and that the sound levels downstream of the duct outlet are higher than expected from using an acoustic monopole radiation pattern. The dominant mechanism affecting the propagation of sound is the refraction arising from the plume's temperature and velocity gradients. The sound propagation is highly dependent on the proximity from the duct outlet, normalised wavenumber, temperature and the jet to cross-flow mean velocity ratio. This computational study builds upon previous experimental work to analyse the fluid-acoustic interaction for heated jets in cooler cross-flow to understand the complex radiation pattern that leads to higher-than-expected sound levels downstream of the duct outlet.

Calibration of Turbulent Model Constants Based on Experimental Data Assimilation: Numerical Prediction of Subsonic Jet Flow Characteristics

Experimental measurements and numerical simulations are two primary methods for studying turbulence. However, these methods often struggle to balance the accuracy and breadth of results. In order to accurately predict the flow characteristics of subsonic jet exhaust and provide a research foundation for the runway crossing operation after the takeoff point, this study utilizes the ensemble Kalman filter algorithm to recalibrate the SA turbulence model constants by integrating NASA’s experimental particle image velocimetry (PIV) data with a sample library generated using Latin hypercube sampling to obtain corresponding flow field calculations. The modified model constants effectively improve the prediction of jet flow characteristics, reducing the spatially averaged relative error along the horizontal axis behind the nozzle from 13.04% to 4.6%. This study focuses on enhancing the accuracy of numerical predictions for subsonic jet flows via the adjustment of turbulence model constants. The recalibrated model constants are then validated to improve the prediction of jet flows under various conditions. The findings have important implications for acquiring high-fidelity data on rear engine jet flows after takeoff, enabling precise determination of safety separation distances, and enhancing the operational efficiency of airports.

Velocity Mapping of an H2 − O2 Exhaust Jet in Air by Means of Schlieren Image Velocimetry (SIV)

Visualization methods have always been used to inspect flows that are invisible to the naked eye. Seedless velocimetry has been regarded as an alternative to other intrusive quantitative methods and adapted to fit many applications in the industrial or scientific field. Schlieren image velocimetry (SIV) uses the general working principle of a schlieren system to acquire flow images, while relying on a particle image velocimetry (PIV)-like algorithm to obtain quantitative data related to the studied flow. The test case of this study consists of a turbulent round exhaust jet generated by a micro-thruster that uses H2−O2 as a propellent. Mapping the local velocities of the flow is achieved by initially performing a lagrangian tracking method which makes use of a direct image correlation algorithm. These results are then compared to the velocity map obtained from a kymograph applied to a series of images. The velocity profiles obtained through SIV will be compared to the velocity profile of the jet provided by the CFD simulation. The schlieren investigation of the jet’s local velocity map is set to determine the thruster’s capabilities, and conclude if the thruster reaches the desired Mach for which it has been designed.

Mathematical Modelling and Fluidic Thrust Vectoring Control of a Delta Wing UAV

Pitch control of an unmanned aerial vehicle (UAV) using fluidic thrust vectoring (FTV) is a relatively novel technique requiring no moving control surfaces, such as elevators. In this paper, the authors’ previous work on the characterization of a static co-flow FTV rig is further validated using the free to pitch dynamic test bench. The deflection of a primary jet after injection of a high-velocity secondary jet was captured using the tuft flow visualization technique, along with the experimental recording of subsequent normal force impinged on the Coanda surface resulting in the pitching moment. The effect of primary and secondary flow velocities on exhaust jet deflection, as well as on the pitch angle of the aircraft, is examined. Aerodynamic gain as well as the inertia of a delta wing UAV test bench are computed through experiments and fed into the equation of motion (e.o.m). The e.o.m developed aided in the design of a model-based PID controller for pitch motion control using the multi-parameter root locus technique. The root locus tuned controller serves as a benchmark controller for performance evaluation of the genetic algorithm (GA) and particle swarm optimization (PSO) tuned controllers. Furthermore, the frequency domain metric of gain and phase margins were also employed to reach a robust control design. Experiments conducted in a simulation environment reveal that PSO-PID results in a better response of the UAV in comparison to the baseline pitch controller.

Improvement of arrangement for including water-jet ejector in CHP operation cycle

The question of the reliability of the operation of the exhaust devices almost always remains relevant during the operation of vacuum deaeration plants, since incorrect operation of the exhaust device entails violations of the vacuum deaeration process. As devices for the removal of non-condensing corrosive gases from vacuum deaerators, jet exhaust devices-ejectors are used, in which the kinetic energy of one stream is transferred to another stream by direct contact (mixing). According to the working body, ejectors are divided into water-jet and steam-jet. In this paper, the current scheme of inclusion of a water-jet ejector in the operation cycle of the Ulyanovsk CHP-2 is considered, problems and shortcomings of the scheme are identified, technical devices and their characteristics are described. The results of tests to determine the temperature and flow rate of the source water supplied to the gas separator tank to ensure stable operation of the EV-340 ejector are presented, which showed that the actual flow rate of the source chemically purified water to the gas separator tank is 17 m3/h, and the water temperature in the gas separator tank reaches 37°C, which exceeds the permissible temperature of the working water of the ejector by 7°C. An improved arrangement for the inclusion of a water-jet ejector in the operation cycle of the CHP is proposed, which provides sufficient cooling of the working water of the water-jet ejector by mixing this water with the source water, which is further cooled in the cooling tower. It is estimated that the new improved arrangement gives an economic effect of 2,5 million rubles/year.