Abstract:
A selectable ramjet propulsion system for propelling a rocket or missile includes a gas generator adjacent a booster. A frangible diaphragm is disposed between the gas generator and the booster. The booster and fuel gas generator can be operated in normal sequence, or operated at the same time in order to increase the thrust produced for short-range missions. A logic circuit contained on the rocket or missile determines a time to rupture the frangible diaphragm based on whether or not the distance to the target exceeds a threshold distance.

Description:
CROSS REFERENCE TO RELATED APPLICATION(S) 
     This patent application claims a benefit to the filing date of U.S. Provisional Patent Application Ser. No. 61/442,929, titled “Selectable Ramjet Propulsion System,” that was filed on Feb. 15, 2011. The disclosure of Ser. No. 61/442,929 is incorporated by reference herein in its entirety. 
    
    
     U.S. GOVERNMENT RIGHTS 
     N.A. 
     BACKGROUND 
     Field 
     The disclosure relates to a propulsion system for a missile. More particularly, applying at least a portion of ramjet impulse during the booster phase improves performance for short range missions. 
     Description of the Related Art 
     A ramjet missile requires a conventional rocket boost phase to reach high supersonic speeds, after which the ramjet can operate as designed. The specific impulse, or fuel consumption, of the ramjet phase is 3 to 4 times more efficient than the rocket phase, but the ramjet phase delivers a lower thrust level over an extended period. Because of this, at short launch ranges, the rocket is able to accelerate and intercept a target in less time than the boosted ramjet. For longer range intercepts, the ramjet is able to propel the missile much greater distances. 
     U.S. Pat. No. 8,056,319 B2, titled “Combined Cycle Missile Engine System,” is assigned to the same assignee as the present patent application. The patent discloses a ducted rocket, also called an air augmented rocket, using the airflow induced into the engine to produce more thrust (augmentation) than would be produced by the rocket alone. Augmentation is low until a flight speed of Mach 1.5 is exceeded. This is due to low ram pressure at low supersonic speeds. Above Mach 1.5, the augmentation rises rapidly and can be in excess of 100%, twice the rocket only thrust. The disclosure of U.S. Pat. No. 8,056,319 B2 is incorporated by reference herein in its entirety. 
     BRIEF SUMMARY 
     One embodiment of the disclosure discloses a selectable ramjet propulsion system for propelling a rocket or missile. This system includes a gas generator adjacent a booster. A frangible diaphragm is disposed between the gas generator and the booster. A booster propellant contained within the booster is ignited and a distance to an intended target is determined. A logic circuit contained on the rocket or missile then determines a time to initiate the gas generator fuel supply and rupture the frangible diaphragm based on whether or not the distance to the target exceeds a threshold distance. 
     It is a feature of certain embodiments that this system allows an engine to burn ramjet fuel and booster propellant when a target is present at short range (less than the threshold distance). This enables the boosted ramjet engine to approach rocket only propulsion performance when long range capabilities are not needed. 
     The details of one or more embodiments of the invention are set forth in the accompanying drawings and the description below. Other features, objects and advantages of the invention will be apparent from the description and drawings, and from the claims. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a variable flow ducted rocket as known from the prior art. 
         FIG. 2  illustrates an unchoked ducted rocket in accordance with an embodiment disclosed herein. 
         FIG. 3  graphically compares ramjet to rocket performance as functions of time of flight and range. 
         FIG. 4  graphically compares the performance of the selectable ramjet propulsion system described herein to rocket propulsion system over a launch range of from 1 to 10 nautical miles. 
         FIG. 5  graphically compares the performance of the selectable ramjet propulsion system described herein to the boost/ramjet propulsion system over a launch range of from 1 to 25 nautical miles. 
     
    
    
     Like reference numbers and designations in the various drawings indicated like elements. 
     DETAILED DESCRIPTION 
     Propulsion phases of a missile mission include: 
     Boost phase occurs when a solid rocket is typically used to accelerate the vehicle to low supersonic speeds where the ramjet engine becomes efficient; 
     Transition phase occurs when the vehicle configuration is changed to allow air to enter the combustion chamber, and the fuel and air combustion process is initiated; and 
     Ramjet sustain phase in which thrust is produced by sustained combustion of the fuel/air mixture. 
       FIG. 1  illustrates a variable flow ducted rocket  10  as known from the prior art. The rocket  10  includes a nose cone  12 , warhead  14 , gas generator  16  and combustor  18 . Nose cone  12  typically includes a guidance system and other electronics. Disposed between the gas generator  16  and the combustor  18  is a fuel control valve  20 . During the boost phase, combustor  18  is filled with a propellant  22  and functions as a nozzled or nozzleless booster operating at high pressure (nominally 2000 psia). At propellant  22  burnout, internal port cover  24  and external port cover  26  are opened to allow inlet airflow through the duct  28 . For the ramjet sustain phase, the gas generator  16  is ignited to produce ramjet fuel. The gas generator propellant  30  is typically an end-burning grain. The gaseous product of the gas generator exits via a choked exit  32  with valve  34  controlling the operating pressure and amount of fuel produced. A typical gas generator  16  operating range is 200 psia-2000 psia. A typical combustor  18  operating range is 112 psia-200 psia, dependent on flight altitude and Mach number. 
       FIG. 2  illustrates a portion of an unchoked ducted rocket  40 . Those portions not illustrated or described herein are similar to the prior art rocket  10  of  FIG. 1 . While described in terms of a rocket, all embodiments apply equally well to missiles. Referring back to  FIG. 2 , as in the prior art described above, a nozzled or nozzleless booster  18  operates at high pressure (2000 psia). At burnout, internal  24  and external  26  port covers are opened to allow inlet airflow through the duct  28 . In the ramjet sustain phase, the gas generator  16  is ignited to produce ramjet fuel, a frangible diaphragm  42  is ruptured, and the gas generator  16  operates at low pressure with the exit  44  unchoked. An exemplary diaphragm is constructed to withstand the boost pressure in one direction, yet yield when pressure is introduced from the gas generator side. For an unchoked gas generator the diaphragm is sized to be large enough to allow subsonic flow at the gas generator exit. The gaseous product of the gas generator propellant  30  communicates with the combustor  18  (12 psia-2000 psia) and adapts to changing flight conditions. A center-perforated grain  46  is used to increase burning surface area and mass flowrate since the propellant burning rate in the combustor  18  is low at low pressure. An exemplary center-perforated grain has a composition of binder, oxidizer, and fuel as established in the prior art, and is sized to maximize fuel loading, yet conform to structural and ballistic requirements. The combustor pressure (a function of Mach number, altitude and flight angle) dictates gas generator pressure and fuel flowrate produced—thereby achieving passive flow control. The missile/engine operates in stable equilibrium at a given Mach number as a function of altitude. 
       FIG. 3  graphically compares ramjet to rocket performance as functions of time of flight and range. Because the booster only operates up to a relatively low Mach number required to start the ramjet, and then ramjet thrust is at a much lower level, the ramjet time-to-target typically suffers for inner boundary  48  relative to the solid rocket. 
     Disclosed herein is a method whereby all or part of the ramjet impulse is applied to the boost phase on command to improve inner boundary performance when desired for a short range mission. This embodiment uses the unchoked ducted rocket engine shown in  FIG. 2 . It can be operated as a ramjet as described above. However, if it is being employed against a short range target and more boost impulse is desired, both the booster  18  and the gas generator  16  can be operated simultaneously to produce more thrust. In this case, the gas generator  16  is ignited first, which ruptures the diaphragm  42  and ignites the booster propellant  46 . The gas generator propellant  30  and the booster propellant  46  then burn simultaneously. Due to the high gas generator surface area, the mass flow at the booster pressure is much greater than for typical end-burning grain design. At the transition to ramjet sustain, a portion of the ramjet fuel is already consumed, but the missile will be at a higher Mach number. Less fuel will remain for sustain operation, so this mode would only be employed where the long range capabilities of the ramjet are not needed. 
     As particularly illustrated in Example 2 below, there is a threshold distance to target beyond which the boost/ramjet mode is superior to the selectable ramjet propulsion system described herein. Preferably, the missile autopilot has access to range information prior to launch and employs logic to select one mode or the other without any input required from the pilot. 
     Advantages and Disadvantages of this Embodiment are: 
     Advantages: 
     Ramjet booster can be sized for takeover only, not increased to improve inner boundary. Increased thrust when desired for short range. 
     Improved ramjet thrust and Mach number, but less fuel remaining at takeover. 
     Disadvantages: 
     Fuel rich exhaust when burning simultaneously, potentially with accompanying unburned carbon. 
     Nozzle sizing and MEOP (maximum expected operating pressure) driven by simultaneous operation. This may result in non-optimum nozzled booster (nozzleless booster may not be affected). 
     The benefits of the preceding embodiments will be more apparent from the Examples that follow. 
     EXAMPLES 
     Example 1 
       FIG. 4  graphically compares the performance of the selectable ramjet propulsion system (reference line  50 ) described herein to rocket propulsion (reference line  52 ) over a launch range of from 1 to 10 nautical miles. This is a computer simulation of an air-to-air engagement. When the enhanced selectable ramjet propulsion boost mode is used, the time to target is better than a comparable rocket within 2 nautical miles. From 4 nautical miles to 7 nautical miles, the selectable ramjet propulsion is within 2% of the rocket time. Beyond 7 nautical miles, the selectable ramjet propulsion is superior to the rocket. The ducted rocket line (reference line  54 ) represents the time to target for a conventional boost to ramjet operation (conserving all the ramjet fuel). 
     Example 2 
       FIG. 5  graphically compares the performance of the selectable ramjet propulsion system (reference line  50 ) described herein to the boost/ramjet mode of propulsion (reference line  54 ) over a launch range of from 1 to 25 nautical miles. The selectable ramjet propulsion mode results in short time to target inside a launch range of 18 nautical miles. For a range in excess of 18 nautical miles, the boost/ramjet mode results in a shorter time to target. 
     One or more embodiments of the present invention have been described. Nevertheless, it will be understood that various modifications may be made without departing from the spirit and scope of the invention. For example, an end-burning gas generator configuration could be employed in the selectable manner, and still demonstrate a degree of thrust increase. Accordingly, other embodiments are within the scope of the following claims.