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Project Overview

The Tadpole Propulsion System is a bi-propellant liquid rocket engine designed to facilitate rapid testing of Vertical Takeoff and vertical landing propulsion components and systems, and to complete the first two milestones of the Collegiate Propulsive Lander Challenge (CPLC). The engine must be capable of Thrust Vectoring 7 degrees in all directions and throttling thrust level to 40% of nominal to achieve these milestones. The system targets 550 lbf of thrust during its nominal operation and is expected to have an Isp of 170-200 seconds throughout its throttle range. Utilizing two linear actuators, the engine can be gimbaled for a thrust vector range of 10 degrees in all directions. The Engine’s propellants are Alcohol fuels with Liquid Oxygen oxidizer. The first revision of the Tadpole Propulsion System will use Isopropyl Alcohol.

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The system is a test-bed engine meant for use on a fixed test stand for development of lander vehicle propulsion system. The general architecture is representative of a lander vehicle propulsion system and the goal is to rapidly test and refine components for a flight engine.

An internal goal for this project is to do more operations and liquid rocket testing in Purdue space program. Previously the testing within the club has been limited to rocket full stacks that happen every ~2 years coinciding with the design cycle of a liquid rocket.

Engine Overview

The engine is the assembly that achieves the thrust output of the system. It consists of the following components: Chamber, Injector, and Igniter. The design emphasizes rapid testing and configuration of these components in development the lander flight engine.

The tadpole engine architecture is capable of testing different fuels between the alcohols (IPA, Methanol, Ethanol), throttle control methods, cooling method configurations, and igniter configurations.

Regenerative cooling on the flight engine is a challenge that the team is anticipating for the flight engine design. Using a water-jacket in the chamber allows for the ability to test the heat-flux through the chamber wall without risk of structural failure. The alcohol fuels are similar, but differing properties between the alcohol propellants offer advantages/disadvantages regarding cooling and engine performance. Rocket engine injector elements are designed to optimize the atomization and mixture of the propellants for high combustion efficiency. This element design will change based on the unique properties of the propellants or the ratio of fuel to oxidizer among other parameters. The tadpole injector features an easily modifiable injector element with a removable pintle that can allow for quick testing of these injector parameters. These features are two examples of how the tadpole engine architecture allows for rapid testing and re-configuration.

Thrust vector and throttle control is another large point of interest that will benefit from testing. Open-loop throttle control will be the first control method that will be tested with this engine in pursuit of the second milestone in the lander challenge. The re-usability of this engine will allow us to then test close-loop control methods. Sufficient testing will be required to optimize the control law parameters for a reliable thrust control.

Designed Performance Parameters

Look at the Tadpole Engine Balance for a complete overview of all tadpole engine performance parameters:

https://docs.google.com/presentation/d/1Y2RYSkp6NNVTPgstSQo9ld-EEuWsrEoLp1QBiWZggd8/edit#slide=id.g28b2d9a278e_0_1

The beige numbers were all parameters chosen during the design of the engine. The other performance parameters are determined by the chamber geometry. Chamber pressure and OF ratio selection are explained below. Exit pressure was chosen using the Kalt-Bendall criterion for flow separation, giving a significant margin at the lowest throttle to avoid flow separation in the nozzle.

Chamber

The chamber is an additively manufactured part with internal channel geometry to allow for coolant flow past the chamber walls. Metal 3D printing was chosen to facilitate complex channel geometry. Short lead times also allow for faster revisions compared to traditional machining methods such as the saddle jacket. By running water through the coolant channels, the tadpole chamber will be used for many tests without structural failure (hopefully). When testing, the pressure and temperature of the water at the inlet and outlet of the water jacket will enable backing out the total heat flux into the coolant during operation, which will be valuable to inform future cooling configuration design for the flight engine. The first tadpole chamber is made out of the Elementum Al6061-Ram2 aluminum alloy and is the first Thrust Chamber Assembly (TCA) to be made out of that material. This material choice was due to a combination of desirable material properties and low cost. The TCA includes the combustion chamber, nozzle, torch igniter interface, gimbal mounts, and coolant channels/manifolding. For more information on the chamber part click here.

Injector

The Tadpole Injector assembly is a static pintle Injector. The selection of the pintle architecture was due to its general excellent stability at low throttle compared to other injector types. Since optimizing for efficiency is not a priority for this design, the lower cstar efficiency associated with the pintle injector is not a major concern. Additionally, the pintle injector is feasibly manufactured in-house on available CNC mills, and PSP has internal design heritage of pintle injectors of which this design is heavily inspired. The injector assembly consists of two 321 Stainless Steel manifold pieces and a removable copper pintle. This allows for the injector element configuration to be modified by replacing the pintle without needing to make new manifolds. The first revision of this assembly is designed for IPA/LOX at a constant OF ratio of 1.2 during its throttle range, and is ox-centered, meaning that the Liquid oxygen is fed into the center pintle piece, injected through the center pintle holes while the alcohol is injected through an annulus surrounding the pintle shaft. For more information on the tadpole injector click here.

Igniter

The Engine is ignited via a Gaseous Hydrogen and Oxygen torch igniter through the side of the chamber. Torch ignition was chosen for its reliability and quick turn around between hot fires on the test stand. The Engine uses a modified Purdue Torch Igniter, which is an igniter that has been used extensively by research groups at Zucrow Labs to ignite a variety of test articles.

The Purdue torch can configured to run at a wide range of mass flow rates and Oxidizer to fuel ratio depending on the application. This is achieved by choosing sonic nozzles upstream of the igniter inlets and picking respective supply pressures to get the desired output. The torch configuration that the tadpole engine will use is not determined yet, and is awaiting post-processing of torch hot fire results.

Integration of the torch igniter presented issues for the tadpole engine. Our early concepts of the injector had a torch igniter going through its face, but issues with sealing and space constraints forced the ignition access to be through the side of the chamber. The side of the chamber, being water cooled, presents its own integration issues since the torch access cannot interfere with the water channels or manifolding. A nozzle was designed to reduce the exit area of the torch flame and to act as a heat sink to protect the chamber wall that is not water cooled. The torch flame will also impinge on the opposite wall of the chamber causing a local hotspot. Due to concerns about melting the chamber wall with the torch flame, the team performed a test campaign to determine the torch configuration that could supply the most ignition energy to the chamber without melting the chamber hardware. Additionally the torch fuel and oxidizer needs to be supplied by flex hoses, since the engine is gimbaling. For more information on the torch igniter integration click here.

Propellants

The tadpole engine utilizes two propellants: liquid oxygen and Isopropyl Alcohol. There was brief consideration of a mono-prop hydrogen peroxide system, which would decrease the complexity and mass of the fluid system of a lander vehicle. The decomposition and handling considerations turned us away from that prospect, although Zucrow Laboratories at Purdue does have experience with the handling of Peroxide and it may be worth re-evaluating for the lander vehicle. The bi-propellant setup will provide better engine performance, and the Alcohol and Liquid Oxygen combination has been used successfully in several Landers Vehicles of similar scope.

Liquid oxygen was chosen as the oxidizer due to its peak performance and operating OF ratio. The low OF ratio is beneficial for regeneratively cooled engine components because it means that there is a larger amount of fuel mass flow rate per total mass flow rate that can be used to cool those components. Liquid Oxygen does have the complications that come with cryogenic fluids, but our club has extensive experience with cryogenics and have successfully used it on our past rockets.

The fuel was narrowed down to kerosene (RP1, Jet-A) and the alcohols (Ethanol, IPA, Methanol) due to their accessibility and compatibility. The kerosene has the highest peak performance and boiling point of the assessed fuels, but its OF ratio that is twice that of the alcohols and its lower specific heat makes it a worse coolant. The cost of RP1 compared to the other fuels, and to a less extent Jet-A, was also a consideration when deciding against kerosene. The alcohols are all great coolants and easy to handle. All three alcohols can be mixed with water to upgrade their cooling properties and decrease combustion temperatures. Isopropyl Alcohol is additionally solvent to PDMS which is a silicon additive that has been successfully used on Masten’s lander vehicles to decrease heat flux while maintaining performance and is potentially a game changer when compared to the other alcohols. We ultimately chose IPA as the fuel for the first revision of the engine, which follows prior art to similar lander vehicles. For a more detailed study comparing the alcohols cooling properties and performance click here.

Chamber Pressure and Oxidizer to Fuel Ratio Choice

Chamber pressure was chosen by trading the performance and feed pressure requirements. Our requirement is to have propellant tank pressures that are less than 1000 psi. By estimating the pressure losses in the fluid system and engine components we can understand the feasibility of chamber pressures given that requirement. (Should update with measured values). The OF ratio was chosen as a compromise between engine performance, combustion temperature and cooling. Lower OF ratio is ideal for regen cooling due to higher coolant mass flow rate and lower combustion temperature, but reduces performance of the engine.

Thrust Vector Overview

The system’s thrust vectoring method is gimbal via two linear actuators and a gimbal ring. The gimbal range is a 10 degree circle from the center. The actuator interfaces with the tadpole chamber which is connected by two brackets that sandwich the mounting holes and the actuator tie-rod. The gimbal assembly and integration was a challenge due to the two propellant feed lines, water inlet and outlet hoses, and the torch igniter feed lines that had to be worked around.

Throttling Overview

Throttle Valves for both propellants, Closed Loop & Open Loop Control, Constant OF ratio. Used quarter turn ball valves, with brushless motor and 1:25 gearbox stack. This was a meant to be a cheap design while also being robust. The motor is controlled by an ODRIVE board. The odrive is commanded via a teensy to perform a certain thrust curve with the valve. The teensy will be capable of reading line pressures and camber pressure for close loovalve. The teensy will be capable of reading line pressures and chamber pressure for close loop throttle control.

Testing

Scope & Requirements

 Milestones