Cold Flow Tests Research Articles

Numerical study of the linear and non-linear damping in an acoustically forced cold-flow test rig with coupled cavities

Quantifying the oscillation amplitude of thermoacoustic instabilities remains a critical and challenging issue, as it is a complex balance between driving and damping processes. The New Pressurized Coupled Cavities (NPCCs) setup designed for the study of acoustic damping is analyzed in this work. It is a cold-flow test rig mimicking the geometry of a liquid rocket engine and equipped with an acoustic forcing device. The chamber 1T mode triggers a strong non-linear harmonic response, while the 1T1L and 1T2L exhibit weak non-linearities. Disturbance energy budgets are used in large-eddy simulations to characterize the damping phenomena with the 1T2L and 1T1L forcing. The correct global damping of the system is retrieved, and local damping contributions are extracted. Then, a non-linear term representing the energy transfer to the harmonics is derived from non-linear acoustics theory. Combined with a linear model, this model correctly retrieves the limit-cycle of the 1T mode.

Effects of slit geometric parameters on spray characteristics of double-slit pintle injectors

Pintle injectors have garnered significant research attention in recent years, particularly for their applicability in reusable launch vehicles, owing to their wide thrust control range and excellent combustion stability. While research has explored the characteristics of pintle injectors in the context of developing these components for actual engine applications, studies focusing on the effects of design parameters on injector performance have been limited. This study investigated the effects of slit geometric parameters, specifically the blockage factor (BF) and slit area ratio (γ), on the spray characteristics of double-slit pintle injectors. Cold-flow tests were conducted using planar pintle injectors with water and ethanol as simulants. The spray angle and Sauter mean diameter (SMD) were measured using the shadowgraph technique, and the distribution of mass flow rate and mixture ratio was analyzed using a mechanical patternator. The experimental results revealed that two distinct streams were injected at different angles from each row of slits, resulting in a division of spray shape, SMD, and mass flow distribution into three regions based on the two streams. These spray angles, termed primary and secondary spray angles, were quantified as functions of the local momentum ratio, determined by BF and γ. To correlate the spray characteristics with combustion performance, mixing quality and a representative droplet size metric, the integral Sauter mean diameter (ID32), were presented. The study found that higher values of BF and γ corresponded to improved mixing quality.

Experimental Study of High-Diodicity Inlet Effects On Rotating Detonation Combustor Performance and Operability

Abstract Implementing high-diodicity inlets is critical to reduce backflow and mitigate the pressure loss across the injector in RDCs. Experiments on air injection diodicity were conducted in a non-premixed RDC for both cold-flow and reacting conditions. The introduction of a Tesla-like diode impacted operating modes and injector dynamics, but the extent of that effect depended on the throat-to-combustor area ratio. A smaller ratio mitigated the impact of the diode on detonation characteristics, while a larger ratio extended the operating range of stronger wave modes. The diode stabilized RDC operation through an increased static pressure drop, but limited performance most likely due to poor reactant mixing and local equivalence ratio distribution. Cold-flow tests showed a higher diodicity for the diode, which may contribute to higher pressure gain in reacting experiments. A modified diodicity formulation based on reacting flow measurements was introduced, suggesting that a multi-metric approach can be useful to assess injector performance. High-diodicity air inlets could be useful tools for reducing total pressure loss and controlling operating modes, but careful consideration is required to limit adverse effects on processes like reactant mixing.

Flame structure of GCH4/GOX pintle model combustor with wall confinement

The pintle injector is a representative option for designing a throttleable rocket engine that can be applied for various applications, including the vertical soft landing of reusable launch vehicles. Although the most direct research method to investigate the characteristics of pintle injectors experimentally is the high-pressure combustion test, it is often limited due to its cost and operational hazards. Consequently, atmospheric combustion or cold flow tests are frequently applied in the rocket injector field to simulate and study the fundamental combustion or spray characteristics. However, previous works, including such simulated experiments under atmospheric conditions, involved an unrealistically wide combustor or open condition. Since only a single pintle injector is typically used per engine, the combustor wall may significantly influence the fundamental characteristics. Thus, in this study, a pintle injector combustor with adequate wall confinement is designed, and its flame structure is investigated across a wide range of total volume flow rates and mixture ratio conditions, incorporating several geometries with varying gaseous oxygen injection areas. This research primarily focuses on the flame structure and its angle. Under specific experimental conditions, the unreported unstable flame was newly observed. The overall experimental results, supported by numerical methods, indicate that wall confinement has substantial importance in simulating a realistic flame structure and internal flow field within the combustor.

Design and analysis of a supersonic diffuser for altitude testing of upper-stage liquid rocket engines

This study conducted a numerical simulation of high-altitude environments for upper-stage liquid rocket engines that use second-throat supersonic diffusers. To obtain the minimum second-throat area, a typical normal shock model was modified and validated using two experimental procedures in an existing high-altitude facility: a cold flow test with a conical nozzle, and a hot test of a 7-ton class liquid engine. The modified normal shock model, known as the passive diffuser model, was employed to design a new second-throat supersonic diffuser and to numerically test it under various conditions. In the numerical analysis, a velocity parameter, such as the Mach number, should be maintained constant with minimal v-velocity at the entrance of the diffuser contraction to realize a good agreement with the quasi-one-dimensional estimation. In the full-scale facility, geometrical parameters of the second-throat supersonic diffuser normally have a marginal effect on the performance of the liquid rocket engine under test, unless they affect the starting condition of the second-throat supersonic diffuser. It is possible to reduce the throttling of the new upper-stage liquid rocket engine to 63 % by utilizing the self-evacuation feature of the newly designed second-throat supersonic diffuser. We find that, as throttling decreases, the hot region shifts to the second throat, necessitating the use of a cooling system to protect the entire second-throat supersonic diffuser from thermal damage.

Experiments and Modelling on the Effect of an Adjustable Boundary on Thermoacoustic Stability

Predicting the existence of thermoacoustic instabilities is a key step in the design of modern gas turbines and the choice of their operating conditions. The stability of a combustion system crucially depends on the acoustic boundary conditions. To systematically investigate the influence of these boundary conditions, a test facility with variable inlet and outlet geometries has been developed. Cold flow tests confirmed that the acoustic terminations allow for a change of the reflection coefficient from close to open end to anechoic to almost closed end over a large frequency range. In the present work, we present the design of an adjustable exit boundary enabling a change in the thermoacoustic stability without modifying the flame operating conditions. Experiments have been conducted in a turbulent axial atmospheric combustor. The acoustic reflection coefficients in hot condition of the exit boundary are measured for different boundary geometries and the impact of these geometries on the flame stability is assessed. A parametrized model is derived and reproduces the experiments well.

Study on Separation Characteristics of Nozzles with Large Expansion Ratio of Solid Rocket Motors

In order to study the flow characteristics of a nozzle with large expansion ratio and its influence on the force on the nozzle, ground cold flow test research and a fluid–structure coupling simulation analysis were carried out (maximum expansion ratio ε = 30.25). The variation and pulsation characteristics of the pressure near the measuring point area and the inlet pressure were obtained through experiments. Through the analysis of the peak-to-peak value and average value, it was found that the average pressure after separation increases by 90%, but the peak-to-peak value increases by about five times, indicating that the pressure fluctuation after separation is much larger than before separation. The separation flow field under cold flow conditions was simulated using the CFD commercial calculation software Fluent to verify the correctness of the numerical calculations. The fluid–structure coupling analysis was carried out on a large expansion ratio (maximum expansion ratio ε = 48) full-scale nozzle, and the structural deformation characteristics of the nozzle under the separation conditions were studied. The research results show that flow separation occurs in the nozzle with a large expansion ratio under ground conditions. Before the separation point, the pressure pulsation on the nozzle wall is small, and the turbulent pulsation effect is weak. After the separation point, the pressure pulsation increases, and the turbulent pulsation effect is enhanced. When the total pressure decreases, the separation area of the nozzle increases, and the separation flow field presents a strong asymmetry. Reducing the total inlet pressure by half resulted in approximately 50 times the lateral load. Under the combined influence of the ground conditions and low total pressure, the large lateral load caused by the asymmetry of the separation flow field will cause the deformation of the nozzle structure to increase by 5.5 times. This research provides an important reference for the design and experiment of nozzles with a large expansion ratio.

Aeromechanical Optimization of Scalloping in Mixed Flow Turbines

Abstract Scalloping of radial and mixed-flow turbocharger turbine rotors has been commonplace for many years as a means of inertia reduction and stress relief. The interest in turbine rotor inertia reduction is driven by transient loading requirements of turbocharged internal combustion engines, as this is a key factor in the time taken to meet transient engine torque requirements. Due to the high-density materials used in turbine rotors, any material removal from the turbine wheel has a significant impact on turbocharger inertia, and thus the transient response of the engine. It is well known that scalloping reduces not only inertia, but also efficiency. This study aimed to identify if it was possible to produce a new scallop design which reduced the scalloping efficiency penalty without increasing inertia or compromising mechanical constraints. This was carried out with the aim of developing design recommendations for scalloping where a complete minimization of inertia is not the design goal. A multipoint, multi-physics numerical optimization, with constraints on inertia and back disc stress, was carried out to determine what efficiency benefit could be realized by aerodynamically designing mixed-flow turbine scalloping. An efficiency benefit was identified across the entire turbocharger operating line, with increased benefit at low engine load, whilst not exceeding the design constraints. Scalloping losses for the baseline design were found to be greatest at low engine load, where the turbine experienced a low expansion ratio, mass flow, and speed. This explains why an aerodynamic redesign yields the greatest benefit under those operating conditions. These performance predictions were experimentally validated on the cold flow test rig at Queen's University Belfast, with good agreement between simulated and measured data. To conclude the study, a detailed loss audit was carried out to identify key loss generating flow structures and to understand how changes in geometry affected the formation and development of these flow structures throughout the passage. A large vortex which entered the passage from the scalloped region and interacted with the tip leakage vortex along the suction surface of the blade was identified as the main source of loss due to scalloping. The optimized design was found to better control the location of entry of this vortex into the blade passage, thus reducing the associated loss and facilitating a performance improvement. Geometric design guidelines were then proposed based on these findings.

Schlieren depiction on flow characteristics and dynamic evolution of RBCC inlet under the combined impact of rocket jet and back pressure

Stable and reliable operation of an RBCC inlet in the extremely wide flight speed range determines the normal operation of the whole RBCC engine. And in different modes, the interior flow characteristics and operational stability of the inlet are impacted comprehensively by both the embedded rocket jet and the back pressure. The flow characteristics and dynamic evolution of a RBCC inlet under the impacts of rocket jet and back pressure are studied by cold flow tests combined with schlieren photography. The experimental results show that: 1) The interior flowfield along the flow passage is differentiated into four distinct zones by the impact of the back pressure, named “easily disturbed zone”, “hardly disturbed zone”, “transition zone”, and “isolated zone”. As the back pressure functions independently, the pressure distribution demonstrates significant stepping characteristic propagating upstream from the “easily disturbed zone” to the “hardly disturbed zone” through the “transition zone”. 2) Low pressure rocket jet relieves the flow separation on the ramp wall, but 150% rise of the back pressure will yield an 80% pressure increase at the isolator exit. The stepping characteristics are weakened during the back pressure develops from the “easily disturbed zone” to the “hardly disturbed zone”. 3) High pressure rocket jet improves the comprehensive back pressure resistance of the RBCC inlet effectively. The pressure at the rocket exit is reduced by 33.45% and the pressure at the isolator exit is decreased by 38.04% when the rocket pressure of Procket/P∞ is increased from 0.95 to 1.91. 4) The stability of the inlet flowfield is synchronously affected by the rocket jet and the back pressure. The independent function of low back pressure at Pb/P∞=0.16 causes obviously low-frequency and wide-range oscillations of the flowfield. As the back pressure is increased by 137.5%, the axial oscillation range is significantly suppressed by 83.1%. The independent function of rocket jet causes no significant oscillation, but as the back pressure of Pb/P∞=0.38 functions collectively, the low pressure rocket jet at Procket/P∞=0.95 drives the interior flowfield return to the oscillation mode of low-frequency and wide-range. Nevertheless, the 105.3% increase of rocket pressure can decrease the axial oscillation range by 7.7%.

Dynamic characteristic modeling and simulation of an aerospike-shaped pintle nozzle for variable thrust of a solid rocket motor

To utilize propulsion energy effectively, it is necessary to develop technology that can enable variable thrust by altering the area of the nozzle throat in a solid rocket motor in real time. In this study, an aerospike-shaped pintle nozzle, an altitude compensation nozzle technology, was designed to improve the performance compared to that of the existing pintle nozzles. Pressure fluctuations caused by the movement of the pintle led to the hysteresis of the combustion chamber pressure, which sharply increased the pintle load. Depending on the design conditions of the actuator, a load that was approximately 2.5 times larger than that in the normal state could be applied to the actuator. Additionally, cold flow tests were conducted under static conditions to verify the reliability of the design code for the actuator load prediction results. Moreover, firing tests were performed to verify the reliability of the internal ballistic simulation code. Finally, factors related to the thermal expansion of the pintle and the deformation of the nozzle throat, which caused changes in the nozzle throat area during combustion, were identified and mathematically modeled. Additional firing tests were conducted to validate the updated numerical model. Before the model was updated, a pressure difference occurred at most three times or more between the performance prediction model and the firing test results, but the performance prediction error was reduced when the modified numerical analysis was used. If the present results are applied to a guided missile system, the high-altitude operation performance will improve and precision thrust control will become possible, enabling flexible guided missile system operation.

Thrust Control of Lab-Scale Hybrid Rocket Motor with Wax-Aluminum Fuel and Air as Oxidizer

This article explores the throttling aspect of the hybrid rocket motor through experiments using a lab-scale motor. The lab-scale motor utilizes a wax-Al based fuel and compressed air as the oxidizer. The oxidizer flow rate was modulated using a PID controller to study the closed-loop thrust control performance of the motor. Numerical simulations and cold flow tests were carried out to identify the suitable gains for the PID control algorithm. Pressure feedback was used in the control algorithm to obtain the closed-loop thrust control. The resultant closed-loop system followed the reference pressure accurately during the step input response test of the system. The maximum error in the observed chamber pressure was 1.86% for a reference pressure of 4.69 bar, which corresponds to a reference thrust of 117.6 N. The response of the system for a ramp input, with linear thrust variation from 78.4 N to 127.4 N, showed that the measured thrust followed the desired ramp profile with a root-mean-square error of 1.99 N. A ramp-down test with the same thrust range produced a root-mean-square error of 6.2 N.

Effect of the recess length on the dynamic characteristics of liquid-centered swirl coaxial injectors

In general, a recess in shear coaxial/swirl coaxial injectors positively affects the stable combustion and combustion efficiency. However, a recess in a liquid-centered gas-liquid swirl coaxial injector may cause spray oscillations, known as self-pulsation. In this study, the effects of the recess length on the spray dynamic characteristics and their mechanisms were investigated. Four liquid-centered liquid-gas swirl coaxial injectors with different recess lengths were manufactured by metal additive manufacturing. Cold-flow tests were performed in a broad range of liquid and gas Reynolds numbers through the injectors. Pressure fluctuations in the gas and liquid injector manifolds were measured using dynamic pressure sensors, and high-speed instantaneous spray images were captured using a high-speed camera. The data obtained from the two measurement methods were analyzed using the fast Fourier transform technique. The recess length affects the frequency and intensity of the self-pulsation and stability map. In all injectors, the self-pulsation frequency increases with the liquid Reynolds number, but the effect of the gas Reynolds number differs with the recess number. Thus, the characteristics of self-pulsation change depending on whether the recess number <1 or ≥1. The present experimental results were utilized in proposing self-pulsation mechanisms based on the recess number.

Injection Properties According to the Inner Shape of Metal Additive Layer Manufactured Coaxial Injectors

The application of three-dimensional (3D) printing technology in rocket engine development has numerous benefits considering cost and time. Among the various techniques of 3D printing, the selective laser melting method has the advantage of being able to manufacture complex structures and process multiple materials. In this study, five types of coaxial injectors with different internal configurations were manufactured using metal 3D printing. To confirm the atomization and mixing performance according to the structure of each injector's oxidizer and fuel post, a cold flow test using water and air was performed under a wide range of experimental conditions. As a result of analyzing the injection pressure drop, discharge coefficient, spray pattern, breakup length, and spray angle, the shape of the oxidizer post had a significant influence on the performance of the injector. In comparison, the effect of the fuel post structure was relatively small; however, there was a meaningful difference in the breakup length and spray angle depending on the direction of rotation.

Real-Time Deep Throttling Tests of a Hydrogen Peroxide Hybrid Rocket Motor

Hybrid rockets based on hydrogen peroxide strengthen the typical advantages of hybrid technology and eliminate its main issues. In fact, when the hydrogen peroxide is decomposed through a catalyst bed, it is possible to design hybrid motors that are simple, affordable, safe, green, storable, reliable, stable, efficient, and compact. Moreover, decomposed hydrogen peroxide is not limited by atomization issues and ignites spontaneously with the fuel, allowing for deep throttling and multiple reignition, pushing hybrids’ energy management capabilities to the top. This can be done simply by controlling the oxidizer mass flow to the combustion chamber. Being an incompressible and noncryogenic liquid, hydrogen peroxide mass flow is relatively easy to control. The authors’ research group has been developing lab-scale hybrid rocket motors that are particularly suited for energy management. A flow control valve has been specifically developed for this purpose. The flow control valve exploits the cavitation on an actuated pintle in order to choke the mass flow. The advantages of such a device, which is called a variable area cavitating venturi, are insensitivity of the mass flow to the downstream pressure variation and linearity of the mass flow with the throat area. The flow control valve has been designed at first, then characterized through static and dynamic cold flow tests. Afterwards, fire tests of a lab-scale hybrid motor have been performed in order to understand the throttling behavior of the complete system. Finally, a real-time control of the oxidizer flow valve, and, consequently, of the hybrid motor thrust, has been implemented and fire tested, showing the impressive on-demand energy management capability of this technology.

Hypergolic Continuous Detonation with Space-Storable Propellants and Additively Manufactured Injector Design

Initiation and sustained detonation of hypergolic liquid rocket propellants have been assessed in an annular combustor. An established impinging injector design was modified for additive manufacturing to improve hydraulic characteristics. Computational fluid dynamics analysis was used to verify decreased flow resistance and increased relative backflow resistance afforded by tapered propellant passageways enabled by additive manufacturing. Two injectors were manufactured and a honing process was employed on one of the injectors to reduce surface roughness associated with additive manufacturing. Cold-flow testing confirmed decreased hydraulic resistance due to tapering of the injector passageways and due to the honing process. The detonability of hypergolic space-storable propellants in an annular combustor with the injectors was assessed experimentally with dinitrogen tetroxide as the oxidizer and monomethylhydrazine as the fuel. The presence of rotating detonation waves was verified via multiple instrumentation methods, including a high-speed camera. The performance of the combustor was seen to increase with the total number of detonation waves, the total propellant flow rate, and the presence of a downstream flow constriction. The work demonstrates continuous detonability of hypergolic space-storable propellants under various operation conditions relevant to in-space rocket propulsion concepts, as well as several benefits of utilizing additive manufacturing for injection-system design.

Experimental investigation of a flow-oriented throttleable injector designed for throttleable hybrid rocket motor

Easily achieving thrust throttling is a significant advantage of hybrid rocket motor. To achieve more stable combustion and reduce the pressure of the upstream feed system during the thrust throttling, it is necessary to control the injection conditions while controlling the oxidizer mass flow rate. In view of this, a flow-oriented throttleable injector that can control the injection pressure drop during the thrust throttling was designed, fabricated and tested. In order to verify the feasibility and obtain the injection pressure drop characteristics, cold flow tests of the flow-oriented throttleable injector were performed, including static injection pressure drop tests with/without back pressure and dynamic injection pressure drop tests with/without back pressure. Then, the steady-state hot test and thrust throttling hot tests of the throttleable hybrid rocket motor with the flow-oriented throttleable injector were conducted, in which the 98%H2O2/PE were adopted as the propellant combination and the controllable cavitating venturi was used to control oxidizer mass flow rate. In thrust throttling tests, the achieved maximum of the thrust throttling ratio was 8.65 under the condition that the change ratio of oxidizer mass flow rate was 7.19. During the thrust throttling, the motor worked stably, and the combustion chamber pressure, thrust and injector pressure drop varied approximately linearly with the linear variation of the oxidizer mas flow rate. The variation of injection pressure drop of the flow-oriented throttleable injector was much less than that of the conventional fixed structure injector, of which the injection pressure drop is proportional to the square of the oxidizer mas flow rate.

Experimental Study on Temperature Change by Cavitation Accompanying Self-Pressurization of Propellant for Small Rocket Engines

Abstract As a propellant for hybrid rocket engines using liquid oxidizer and solid fuel and for liquid rocket engines, the use of self-pressurized fluids such as nitrous oxide has become widespread. Since these fluids can be self-pressurized by their high saturated vapor pressure, the propulsion system becomes smaller and simpler. However, this self-pressurization generally forms a gas–liquid two-phase flow by flashing or cavitation. This flow is considered highly unsteady because the temperature and pressure greatly change with the discharge process. In this study, unsteady flow characteristics due to self-pressurization were experimentally obtained by conducting many cold flow tests with carbon dioxide as self-pressurizing fluids. As a result, it was clarified that the fluid temperature dropped about 10–15 K with the pressure drop due to feedline pressure loss during the discharge process. From these experimental results, we estimated the bubble growth and void fraction change that would satisfy the temperature drop. In this paper, the obtained test results and estimated temperature drop are reported.

Direct Formulation of Bipropellant Thruster Performance for Quantitative Cold-Flow Diagnostic

We present a straightforward formulation predicting the characteristic velocity and specific impulse for bipropellant thrusters as a direct function of injection conditions, propellant combination, and nozzle expansion ratio. The theoretical formulation deduces a framework for a quantitative noncombustion test or cold-flow test to predict the performance indices. We simply employ water and dyed water as simulant liquids and then measure the local ratios of mixture and flow rate using an absorbance spectrometer. Density ratio mismatch between hypergolic propellants and water can be reasonably convertible. Combined with chemical equilibrium analysis, we obtain characteristic velocity and specific impulse across wide injection mixture ratios. The validity of the quantitative water-flow diagnostic is evidenced by comparing the results with those of corresponding combustion tests using nitrogen tetroxide and monomethylhydrazine as the propellants under several injector unlike-doublet and triplet configurations, showing that the mixing states of bipropellant thrusters under combustion can be reproduced using the water-flow diagnostic.

Development and Testing of Liquid Simulants

A group of liquid simulants was developed in order to ease the process of cold-flow testing by lowering associated costs and hazards while providing accurate analysis of desired liquid behavior through an injector. Simulants were made with the intention of modeling certain physical properties of the liquid such as density, viscosity, and surface tension. These simulant solutions are capable of modeling liquid propellants throughout large physical property ranges by simply modifying the ingredient ratios. The effectiveness of these solutions in simulating their intended liquids was measured by performing various density, viscosity, and surface tension experiments along with cold-flow tests comparing the liquid breakup and pressure drops across the injector as a function of Reynolds number, Weber number, Ohnesorge number, and mass flow rate. The experimental tests proved that the simulants tested accurately demonstrated desired liquid density, viscosity, and surface tension values and were more effective at modeling jet breakup and pressure drop behavior of specific propellants than that of traditional cold-flow experiments using de-ionized liquid water, with little-to-no added maintenance or cleanup.

Cold flow testing and CFD analysis of screw pump for LOX pump of semi-cryogenic engine

Cold flow testing and CFD analysis of screw pump for LOX pump of semi-cryogenic engine