AAWSAP DIRD, Detection and High-Resolution Tracking of Vehicles at Hypersonic Velocities, November 2010

This Defense Intelligence Reference Document was prepared by the Defense Intelligence Agency's Defense Warning Office in fiscal year 2010 as part of the Advanced Aerospace Weapons System Applications program. It reviews how air flows around objects at subsonic, supersonic and hypersonic speeds, covering shock waves, wakes and ionization. It then compares ways to detect and track hypersonic vehicles, including radar, optical, infrared, LIDAR, infrasound and seismic methods. The report makes four recommendations, among them building a database of aircraft wake signatures and developing novel detectors.

From the source: Release of 2026-09-18 Incident: 11/20/10, Las Vegas, Nevada. Released with redactions. This document is a Defense Intelligence Reference Document (DIRD), a technical reference format used by the Defense Intelligence Agency (DIA) to capture baseline knowledge on a specific topic for later analytic use. DIRDs are best understood as reference and synthesis products rather than as original research. It is one of 38 DIRDs produced under the Advanced Aerospace Weapon System Applications Program (AAWSAP) between 2009 and 2011. Because AAWSAP’s scope permitted a broad range of supporting topics, not every DIRD in the series directly concerns aerospace systems or future threat assessment. The following summary reflects the DIRD’s scope and framing at the time of writing and should not be read as implying current validation of the concepts discussed. This DIRD surveys how hypersonic vehicles may be detected and tracked by exploiting the physical effects they create in flight, especially shock waves, ionized gas, hot surfaces, and turbulent atmospheric wakes. The report reviews a broad set of detection methods, including radar, infrared sensing, optical imaging, LIDAR, passive radio reflection, infrasound, and seismic techniques, and argues that the most effective systems will likely combine multiple sensor types, because each captures different features of a high-speed vehicle’s passage through the atmosphere. It presents radar and infrared sensing as the strongest existing tools for operational detection, while giving particular attention to wake-based methods such as LIDAR and passive radio techniques for improving tracking, identification, and discrimination of future hypersonic aircraft. The paper identifies hypersonic vehicles as an evolving surveillance problem in which future progress will depend on better multi-sensor integration, improved wake characterization, and novel signature-exploitation techniques.

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These equations can be expressed in a form where the ratio of density, temperature, and
pressure across the normal shock are a function of the upstream Mach number, M1.
(11)
(12)
Pz = 1+ ~(M12 - 1) (13)
P1 r+ 1
Pi <r + t) M 12
= ----- (14)
Pi 2 + (y - 1) M i2
As an example, if the object shown in figure 4 is traveling at Mach 3 through air at sea level,
the pressure in the reg ion between the nose of the object and the shock changes as shown in
the following table:
Table 1: Normal Shock Values for Mach 3 Flow
Freestream Ratio of Values
Values Properties Downstream
Across the of the
Normal Shock Normal Shock
M1 =3 M2/M1 = 0.158 M2 = 0.475
p1 = 101,325 Pa p/p 1 = 10.333 p2 = 1,047,000 Pa
T1 = 293 K Ti/T1 = 2.679 T2 = 785.0 K
= (2o·q = (s12·q
p1 = 1.293 kg/m 3 p/p 1 = 3.857 p 2 = 4.976 kg/m 3
For a normal shock, the Mach number on the downstream side of the shock is always
subsonic, and in th is example, the Mach number abru ptly drops from M1 = 3 to M2 = 0.475.
The temperature always increases across a shock, and for a Mach 3 flow, the temperature
rises from room temperature at 20° C up to 512° C. It is apparent that, for supersonic
aircraft traveling at this speed, the fuselage and wings will need to be made of materials that
can withstand the high temperatures and the dramatic pressure increase of 10.333. Shocks
are always accompanied by significant increases in pressure, temperature, and density in the
flow across the shock.
Flow about an Ogive or wedge-shaped airfoil as shown in Figure 4 can be analyzed in a
similar fashion; however, the wall deflection angle, e, will affect the flow and the formation of
the oblique shock wave that surrounds the airfoil.
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