3D computational study of thrust vectoring using bypass mass injection in a propulsion nozzle

The transverse injection in a supersonic cross-flow is problematic which can be encountered in several aerodynamic applications such as fuel injection in scramjet combustor, missile control, drag reduction, and thrust vector control. In recent years, an extended analytical, numerical, and experimental work has been carried out by the authors [1, 2] to investigate the vectoring performances of a supersonic axisymmetric nozzle using secondary fluid injection. Secondary gas injection thrust vector control (SITVC) or shock vector control (SVC) is considered as an alternative way to control the thrust direction of a rocket nozzle beside the classical use of mechanical device such as fluidic actuators. In the context of SITVC operation, the nature and source of injectant gas may raise efficiency-related issues. In previous studies [3–6], it is well established that injection of gas with low molar mass promotes better jet penetration and therefore will be a better choice for SITVC operation. To assess this point, an experimental test campaign has been conducted in the hypersonic test facility EDITH of the CNRS institute ICARE in Orleans, France. The focus of the study is to analyze the secondary injectant gas thermodynamic properties influence on the global vectoring performance of a supersonic nozzle. For this purpose, performance aspects of fluidic thrust vectoring concept have been experimentally investigated on a truncated ideal contour (TIC) nozzle model using a variety of gas species (with low to moderate molar mass) as injectant. Qualitative and quantitative diagnostics consisted of Z-Schlieren visualization, 3-axis force balance, and static and dynamic parietal pressure measurements. The experimental results are compared to the numerical and analytical findings.

In recent decades, the fluidic thrust vector control technique is one of the core strategies to redirect various aerospace vehicles, such as modern launch rockets, supersonic aircraft, and guided missiles. The fundamental theory of the shock vector control is that the gas is injected into the supersonic part of a conventional convergent–divergent nozzle from the transverse to cause interactions between the shock waves and boundary layers. Then, the deflection of the primary jet can be easily realized by the induced oblique shock waves. It was evident that the shock vector control is a very simple, low weight, low cost, and quick vectoring response technique to gain high thrust vectoring performance. In the present work, computational fluid dynamics studies were performed at different control parameters in a three-dimensional rectangular supersonic nozzle with the slot injector. For the validation of the numerical methodology, computational results were compared with experimental data referred to the NASA Langley Research Center. The static pressure distributions along the upper and lower nozzle surfaces in the symmetry plane were matched with the test data excellently. Numerical simulations were based on the well-assessed shear stress transport k–ω turbulence model. Second-order accuracy was selected to reveal more details of the flow-field as much as possible. Lots of factors were investigated, such as the momentum flux ratio, length-to-width ratio, injection location, and injection angle. The performance variations for different affecting factors were illustrated and some constructive conclusions were obtained to provide the reference for further investigations in fluidic thrust vector control field.

Thrust Vectoring Applied to Catastrophic Failure Prevention in Jet Transports

Modelling and numerical analysis of aluminium SITVC C-D nozzle for thrust vector control

Thrust vector control technology is widely used in modern advanced aircrafts, space shuttles and unmanned aerial vehicles, etc. An accurate thrust vector mathematical model is of great significance to enhance the thrust vector control effect. In this paper, the P200 micro-turbine engine with an axisymmetric thrust vector nozzle placed at the rear is taken as an example, A hybrid thrust vectoring dynamics modeling method consisting of engine speed model, engine thrust model, actuator dynamics model, airflow angle model and thrust loss coefficients is proposed. The engine speed model is of Linear Parameter Varying (LPV) model with each segment is characterized by a First Oder Plus Dead Time (FOPDT) model. The engine thrust model is characterized as a cubic polynomial function of engine speed. The actuator’s dynamic model is a FOPDT model. The nozzle airflow deflection angle is proportional to the airflow deflection angle. the thrust loss coefficient is approximated by a cosine relationship with the airflow deflection angle. The experimental results show that the established model can well characterize the dynamics of the thrust vector system.

In this research, the effects of some liquid side injection from nozzle wall into exhaust gas of combustion chamber are studied. The side injection against main flow is as elliptical solid thing that change the symmetric of flow field on nozzle wall and causes some different pressure distribution on wall, and finally causes thrust vector deviation. Flows interaction causes some physical phenomena as bow shock wave in front of injection region. This paper explain the effects of this wave and variation velocity & pressure distribution at different cross sections of flow field and comparison results of air and other liquid fluid in thrust vector control system. The results are compared with experimental data and have well agreement with them. The results show that Freon is one of best injection liquid for this type of thrust vector control. Performance of Injection is optimum in relative position 35 to 40% nozzle divergence length.

A potential concept that can offer an alternate method for thrust vector control of the Space Shuttle Solid Rocket Booster is the use of a cylindrical probe that is inserted (on demand) through the wall of the rocket nozzle. This Probe Thrust Vector Control (PTVC) concept is an alternate to that of a gimbaled nozzle or a Liquid Injection Thrust Vector (LITVC) system. The viability of the PTVC concept can be assessed either experimentally and/or with the use of CFD. A purely experimental assessment can be time consuming and expensive, whereas a CFD assessment can be very time- and cost-effective. Two key requirements of the proposed concept are PTVC vectoring performance and the active cooling requirements for the probe to maintain its thermal and structural integrity. An active thermal cooling method is the injection of coolant around the pheriphery of the probe. How much coolant is required and how this coolant distributes itself in the flow field is of major concern. The objective of the work reported here is the use of CFD to answer these question and in the design of test hardware to substantiate the results of the CFD predictions.

Methods of vectoring and modulating exhaust thrust in a converging-diverging nozzle by secondary fluidic injection were investigated. The application of fluidic thrust control (FTC) offers potentially significant gains in performance and manoeuvrability without the cost of heavy mechanical systems. FTC nozzles also have significant advantages in relation to reducing observability and are particularly suited to low-cost, lightweight, highly manoeuvrable missiles and uninhabited combat air vehicles. The throat shifting thrust modulation (TSTM) and the shock vector control (SVC) FTC methods, and the combination of the TSTM and the SVC systems to achieve both modulation and vectoring simultaneously, were investigated to determine the level of flow interaction and any resulting influence on thrust. In the SVC method, shocks were induced in the supersonic flow by secondary injection in the diverging section of the nozzle. In the TSTM method secondary fluid was injected near the nozzle throat. A design of experiments (DOE) approach was used to study the effects of different factors on the fluidic thrust control. In the TSTM method study, relative injector size (Aratio), injector angle (θinj) and relative injection pressure (Prratio) and the interaction of the latter two factors were found to be significant to the effectiveness of the thrust modulation. In the SVC method the maximum thrust vectoring angle was achieved for the case where the induced oblique shock was just impinging at the opposite side wall at the nozzle exit. The throat modulation efficiency reduced for the combined method due to the additional SVC injection mass flow. The same amount of vectoring was achieved with a lower injection mass flow for the combination method compared to the standalone SVC method.

An experimental assessment of thrust vectoring of an overexpanded, circular supersonic jet is performed by using an array of 64 individually addressable microjet-in-crossflow actuators located at four distinct streamwise locations. A genetic algorithm, using the pressure distribution measured by a rake of total pressure probes in front of the jet and having the goal of maximizing the thrust vectoring angle, achieved a maximum vectoring angle of 2 deg after exploring 900 distinct actuator configurations. The solution is also demonstrated to provide azimuthal control of the vectoring direction due to the circular cross section of the nozzle, which is useful for rocket control systems. Further analysis of the optimal solution indicates the genetic algorithm leveraged the production of an asymmetric shock structure at the diverging section of the nozzle to achieve shock vector control.

A parametric investigation of Dual-throat nozzle bypass channel configurations for advanced Aviation applications

Thrust vector control (TVC) might be used to control aircraft at large altitudes and in post-stall conditions when aerodynamic control surfaces are ineffective. This study demonstrated that the implementation of the TVC on high-speed aircraft is a reasonable solution and might be an alternative when compared to the complicated reaction control system or large aerodynamic control surfaces. The numerical flight dynamics model of the X-15 experimental aircraft was developed and implemented in MATLAB/Simulink and then used to investigate the proposed solution. The obtained results indicate that the aircraft, equipped with full 3D thrust vectoring and two independent horizontal stabilizers to control the roll angle, was able to achieve flight along the path that was defined by a set of waypoints. This paper also highlights the potential benefits and challenges of using TVC as a control method for aircraft. The results of this study contribute to the growing body of research on aircraft control and simulation. Future work can explore the use of TVC for other aircraft with unique configurations and low maneuverability features.

Thrust vector control (TVC) methods were developed as an alternative to aerodynamic control surfaces for missiles, which suffered from poor pitch-over characteristics in launching stage. TVC system, can allow the missile to turn quickly and stably in any direction by deflecting the nozzle exhaust gas irrespective of the missile speed. A jet vane is one of such devices, used for thrust vectoring. In conventional approaches, jet vane is connected control surface through a mechanical interface such as wire and link. However, these approaches lead to angular position dependent transmission ratio. In addition, these techniques limit designer to obtain the required hybrid control ratio. In this study, a novel compact electromechanical actuator is proposed to overcome these problems. Constant or variable transmission ratio can be obtained using such actuator. Proposed solution is not only a simple and compact structure, but also being able to provide sufficient control torque for a missile's quick turn.

Two controllers are developed to provide attitude control of a spinning rocket that has a thrust vectoring capability. The first controller has a single-input/single-output design that ignores the gyroscopic coupling between the control channels. The second controller has a multi-input/multi-output structure that is specifically intended to account for the gyroscopic coupling effects. A performance comparison between the two approached is conducted for a range of roll rates. Each controller is tested for the ability to track step commands, and for the amount of coupling impurity. Both controllers are developed via a linear-quadratic-regulator synthesis procedure, which is motivated by the multi-input/multi-output nature of second controller. Time responses and a singular value analysis are used to evaluate controller performance. This paper describes the development and comparison of two controllers that are designed to provide attitude control of a spinning rocket that is equipped with thrust vector control. 12 refs., 13 figs., 2 tabs.

The experimental article focuses on the flow characteristics and acoustic features of a jet issuing from a Convergent Divergent (C-D) nozzle with a cruciform (cross-shaped) cross-section. The design Mach number is 1.5. The cruciform nozzle has an equivalent exit diameter of 5.42 mm, and its features are compared with the equivalent circular and square nozzles. The operating Nozzle Pressure Ratio (NPR) is varied from 2 to 6. The pitot probe and microphone, in conjunction with data acquisition instruments, provide the total pressure and acoustic measurements of the jet flow. The cruciform jet exhibits a shorter supersonic core length compared to both the circular and square jets for all operating NPRs. Compared to circular and square jets, the cruciform jet has a maximum Overall Sound Pressure Level (OASPL) reduction of 7.06 dB and 7.73 dB at NPR 5. The screech is suppressed and is only discernible at NPR 5 in the cruciform jets, whereas it exists from NPR 3 to 6 in the other two jets. The higher screech amplitudes are localized to lower emission angles from 45° to 75°. Together, the OASPL, and screech are suppressed in cruciform jets.

Future space missions may use Nuclear Thermal Rocket (NTR) stages for human and cargo missions to Mars and other destinations. The vehicles are likely to require engine thrust vector control (TVC) to maintain desired flight trajectories. This paper explores requirements and concepts for TVC systems for representative NTR missions. Requirements for TVC systems were derived using 6 degree-of-freedom models of NTR vehicles. Various flight scenarios were evaluated to determine vehicle attitude control needs and to determine the applicability of TVC. Outputs from the models yielded key characteristics including engine gimbal angles, gimbal rates and gimbal actuator power. Additional factors such as engine thrust variability and engine thrust alignment errors were examined for impacts to gimbal requirements. Various technologies are surveyed for TVC systems for the NTR applications. A key factor in technology selection is the unique radiation environment present in NTR stages. Other considerations including mission duration and thermal environments influence the selection of optimal TVC technologies. Candidate technologies are compared to see which technologies, or combinations of technologies best fit the requirements for selected NTR missions. Representative TVC systems are proposed and key properties such as mass and power requirements are defined. The outputs from this effort can be used to refine NTR system sizing models, providing higher fidelity definition for TVC systems for future studies.