Sıvı Yakıtlı Roket Motorları İçin Eş Merkezli Girdap Enjektörü Tasarımı

Abstract

Sıvı yakıtlı roket motorlarında yakıt atomizasyonun istenilen karakteristikte ve performansta olması; roket motorunun beklenen özgül itki değerlerini yakalaması açısından hayati önem teşkil etmektedir. Literatürde yapılan incelemelerde sıvı yakıtlı roket motorları için geliştirilen ilk enjektörlerde Rus tasarımı ürünlerin daha çok analitik ifadelerin türetilmesi ile belirlendiği, ABD tasarımı ürünlerde ise deneysel veriler kullanılarak türetilen ampirik bağıntılar kullanılarak atömizör tasarımları yapıldığı görülmüştür. Yapılan ilerlemeler ile analitik ve ampirik bağıntıların geliştirildiği, var olan sonuçların nümerik çalışmalarla da simule edilerek doğrulandığı görülmüştür. Sıvı yakıtlı roket motorlarında kullanılan yakıtlar genellikle kriyojenik özelliktedir ve yanmanın gerçekleşmesi için sıvı kütlenin hızlıca buhar fazına geçirilmesi bir gerekliliktir. Bu sebepten yanma odası girişinde yakıtlar enjektörlerde farklı metodlar kullanılarak küçük taneciklere ayrışması sağlanmaktadır. Küçük taneciklere ayrışan sıvının moleküllerinin yüzey alanı ciddi oranda artmaktadır ve buharlaşması için gerekli olan enerji ihtiyacı düşmektedir. Buharlaşma için gerekli olan enerji ihtiyacının azalması, sıvının buharlaşma hızını arttırmaktadır. Buhar fazına geçen yakıt (yanıcı+yakıcı) bir enerji kaynağı ile reaksiyona girmesi sağlanır ve yanma başlar. Burada kritik olan buharlaşmanın enjektör çıkışından sonra hangi mesafede en yoğun şekilde gerçekleşeceğinin belirlenmesidir. Çünkü ilgili mesafe, aynı zamanda yanmanın nerede başlayacağını, dolayısı ile motor performansını direkt etkilemektedir. Bu sebepten sıvı yakıtlı motor yanma odası tasarlanıyorken, enjektörün ortalama tanecik çapı, sprey koniklik açısı ve tanecik parçalanma boyununda tasarıma girdi parametre olarak göz önünde tutulması gerekmektedir. Yapılan çalışmada, atomizör için ister olarak belirlenen parametreler; verilen girdi parametreleri kullanılarak hem analitik olarak türetilmiş hem de nümerik olarak doğrulanmıştır. Girdi parametresi olarak verilen yaktı debisi ve basınç düşüm değeri kullanılarak ortalama tanecik çapı, sprey koniklik açısı ve tanecik parçalanma boyu analitik olarak belirlenmiştir. Yapılan analitik hesaplamalarda viskoz olmayan akış şartları göz önünde bulundurularak ilgili denklemler türetilmiştir. Denklemler türetilirken Bernolli eşitliği, kütle ve enerji korunum denklemleri ile açısal momentum korunum denklemleri kullanılmıştır. Analitik olarak isterlere uygun olarak boyutlandırılan ve tasarlanan atomizör için GAMBIT kullanılarak kare çözüm ağı oluşturulmuştur. Oluşturulan çözüm ağı Fluentte çözdürülmüştür. Yapılan akış analizinin gerçeğe en yakın olabilmesi için iki fazlı akış çözümü zamana bağlı olarak çözdürülmüş ve başarılı şekilde iç akışının oturduğu şartlar belirlenmiştir. Geometrik boyut, çözüm ağı çözünürlüğü ve akış hızları göz önünde bulundurulduğunda zamana bağlı çözümde kullanılan ilerleme adımlarının mikro saniyeler mertebesinde olduğu yapılan çalışmalarla görülmüştür.

Thrust request of aerial vehicles is provided from different type of propulsion systems for atmospheric and exoatmospheric application. In atmospheric operations, air breathing or propeller engine systems could be used for a propulsion request; however rocket engines are the only choice for exoatmospheric operations since there is no air in that environment. Rocket engines could be classified into the two main types: One is the solid rocket engine and the other is liquid rocket engine. In liquid rocket engines (LRE), both fuel and oxidizers are stored in liquid phases and general configuration includes tanks, pump systems, valve system, injection systems, combustion chamber, nozzle and cooling passage for LRE’s. Liquid fuel and oxidizers are both pressurized with pumps and transferred to the injection system. After atomization starts, the following step is the vaporization of liquid components. Finally, combustion takes place from fuel and oxidizers vapor mix. Both liquid and solid rockets have some advantages and disadvantages. Liquid rocket engines have higher thrust level, led to adjust the thrust level and have stop-restarting features. Besides these advantages, liquid rocket engines have also some disadvantages. Cryogenic liquid storage, complex flow control systems and high maintenance costs are the basic disadvantages of this type of engines. Conversely, solid rocket engines have less complex systems and this gives an advantages to solid engines such as storing propellant is simple and firing system of propellant is simple. Adjustable thrust level and restarting is not applicable for solid rocket engines. Liquid rocket engines are used most of space shuttle and launching vehicles as the main thrust system. Combustion performance of liquid engines has directly effect on total motor performance. Key issue for high combustion efficiency is to fine atomization and mixing of propellants. Hence, atomization quality and stability of injection system has a critical role on the total performance of rocket engine. Liquid atomization is the process that is the disintegration of bulk liquid into the small droplets. Liquid atomization is widely used in different industries like thrust chambers, pharmaceutical industry and agriculture industry for different aims. Obtaining the expected design characteristics and performance parameters of injectors in liquid-propellant rocket engine have critical role on rocket engine performance output. Russian researchers had a first studies on coaxial injectors to express the phenomena inside the atomizer by deriving analytical expressions. Besides researchers from US expressed the empirical equations derived from experimental data to clarify the inside flow of an atomizer. Progress in both analytical and empirical studies resulting more precise and controlled injector design. Liquid-propellant rocket engines mostly use cryogenic propellants and the rapid vaporization of liquid bulk is a necessity for high performance combustion in thrust chamber. Hence, different types of injectors could be used according to the requests in combustion chamber to disintegrate the bulk liquid propellant. After disintegration of bulk liquid, droplets have much higher surface area than bulk liquid and this led them a lower energy needs to vaporization. Decreasing requested energy for vaporization of droplets led to increase the vaporization speed. Combustion starts with igniting when the both propellant are in vapor phase. Critical parameter for the combustion is the distance between the injector and vaporization zone since this distance says where the combustion starts. Therefore, injector droplet mean diameter, spray cone angle and break-up length parameters take into consideration when designing a combustion chamber of liquid-propellant rocket motor. In liquid rocket engines, there are different types of injectors, which could be used. Coaxial and impinging types are two main classification and both of them have some advantages and disadvantages. Coaxial types of injectors are widely used especially in liquid oxygen and liquid hydrogen propellant engines. Coaxial injectors are can be divided two sub-group: Shear Coaxial and Swirl Coaxial. Swirl coaxial types of injectors have two swirl type injectors located coaxially. Inner swirl injector atomizes the liquid oxygen and the outer swirl injector atomizes the fuel. Coaxial swirl injector can classified as a sub-type of swirl atomizers. These type atomizes have been started to develop since beginning of twentieth century. Studies on this type injector were mostly focused on the inner flow phenomena of atomizer at early times. Analytical equations were derived for the swirl atomizers with the assumption of inviscid liquid. Derived equations were improved by using the experimental data and could be useful for swirl atomizers with the different geometries. Droplet size, spray cone angle, flow rate, and most of other important parameter effect the flow could be determined by the derived equations. Swirl atomizers used in rocket engines have some basic characteristic than used in other application like gas turbine. Higher flow rate and cryogenic flow condition in rocket engines are the main differences than the other applications. This situation led the designer to use a different design procedure especially on materials and mechanical sides. Nowadays numeric analysis used in computer led to designer to characterize the atomizer rapidly and the modern test facilities are give valuable characterization data to the designer. Swirl atomizers could be named as the simplex atomizer or pressure-swirl atomizer literally. Working principle of swirl atomizer is quite simple; however, hydrodynamic and aerodynamic flow phenomena of swirl atomizer are highly complex. In swirl atomizer, fluid is entering swirl chamber via the tangential port to increase the swirl velocity component. After entering swirl chamber, fluid is directed in to the conical surface to reach the exit orifice. Highly swirled fluid exits atomizer with the high relative speed to the stationary gas phase into the atomizer environment and this led to fluid to starting the disintegration. In this study, pre-determined performance output of the injector are find out according to giving input parameters via the analytical and empirical equation, and numeric validation. Injector input parameters (propellant mass flow rate and pressure dropt) are used in analytical equations to calculate the internal geometry of injector and spray performance characteristics (mean droplet diameter, spray cone angle and break-up length). Analytical equations are derived according to the inviscid theory and derived by using the Bernolli, mass, energy and angular momentum conservation equations. GAMBIT mesh generator program is used in this study as a part of numeric works to create the mesh in the interior face of injector. Numerical analysis are performed in Fluent and for an injector realistic analysis, numerical solution is solved transiently in two-phase flow. Geometric dimensions, mesh geometry and fluid velocity are directly affecting the convergence criteria, and time steps for injector numerical analysis are found around the microseconds.